Interception missle and warhead therefor

ABSTRACT

A fragmentation warhead is provided, capable of being mounted in a carrier vehicle, the warhead having a longitudinal axis. In at least one example the warhead includes a shell that extends along the longitudinal axis. The shell includes a fixed shell portion and a fragmentation portion, and defines therebetween a cavity for accommodating therein an explosive charge. The fragmentation portion includes at least one set of serially adjacent fragments in correspondingly serially contiguous relationship in the fragmentation portion and in generally helical relationship with respect to the longitudinal axis. A corresponding carrier vehicle and a corresponding missile are also provided.

TECHNOLOGICAL FIELD

The presently disclosed subject matter relates to interception missilesand warheads therefor, and to systems and methods using such missiles.

PRIOR ART

References considered to be relevant as background to the presentlydisclosed subject matter are listed below:

-   WO 2013/105093-   WO 2006/238403-   U.S. Pat. No. 4,093,153-   U.S. Pat. No. 7,977,614-   U.S. Pat. No. 6,209,820-   WO 2010/125569-   US 2005/077424

Acknowledgement of the above references herein is not to be inferred asmeaning that these are in any way relevant to the patentability of thepresently disclosed subject matter.

BACKGROUND

Ground to Ground (GTG) rockets such as the GRAD are a serious threatwhen hitting a civil populated territory.

The simple conventional GTG rocket structure has a clear inherentadvantage of very simple launching means and a low-cost price tag,facilitating deployment of numerous launchers and utilization ofnumerous rockets which can be launched in large numbers and over a longperiod of time towards a desired territory, causing not only damage inproperty, injury and loss of life, but also substantial damage to themorale of the population which is subjected to prolonged and continuousthreats.

Recently a new anti-rocket system called the “IRON DOME” system(provided by RAFAEL, Israel) has been introduced and used by the IsraelDefense Forces (IDF) for protecting various towns that have beentargeted by GRAD and other rockets, launched from the Gaza Strip.

A significant challenge in the interception of GTG rockets relates tothe lethality problem: the vulnerable area of the rocket has relativelysmall dimensions and is surrounded by one or two layers of steel-madeballs fitted onto the warhead explosive. Achievement of lethalinterception is possible by using the hit-to-kill technique or byhitting of the threat's warhead by the very dense beam of relativelyheavy fragments (e.g. known per se tungsten fragments of at least 35-40grams).

Usage of hit-to-kill techniques for interception of spinning rocketsappears to be problematic due to precession movement of the rocket withunpredictable amplitudes. On the other hand, the killing mechanism basedon a fragmentation warhead requires a relatively large number of heavyfragments accelerated to required velocity to achieve the lethal effect.Dimensions and weight of the warhead directly influence the size,weight, and cost of the interception missile. The number of requiredlethal fragments generally depends on the volume of uncertaintyregarding the position of the threat relative to the interceptionmissile during the end game. The conventional approach for reducing thisuncertainty volume is by using different on-board sensors such as forexample RF seekers or electro-optical sensors. Another approach forreducing the uncertainty volume is disclosed in WO 2013/105093. Thispublication, by way of general background, discloses a counter-flyingobject system comprising a sensor array including at least one activesensor and at least two passive sensors configured to detect and trackthe flying object, and a missile launcher configured to launch aninterceptor to intercept the flying object, wherein upon launching theinterceptor, the sensor array is configured to determine the location ofthe interceptor and the flying object and send said object andinterceptor locations to a control system, the control system beingconfigured to provide mission data to the interceptor based on saidobject and interceptor locations for guiding the interceptor toward theflying object and activating a fragmentation warhead on or in thevicinity of the flying object when a lethality criteria is met.

Conventionally, such constraints compel utilization of a sophisticatedand costly defense system, for defending against a comparatively simpleand low-cost rocket launch system.

The contents of the above references are incorporated herein in theirentirety.

General Description

According to a first aspect of the presently disclosed subject matter,there is provided a fragmentation warhead configured for being mountedin a carrier vehicle, the warhead having a longitudinal axis andcomprising:

-   -   a shell extending along said longitudinal axis and comprising a        fixed shell portion and a fragmentation portion, and defining        therebetween a cavity for accommodating therein an explosive        charge;    -   the fragmentation portion comprising at least one set of        serially adjacent fragments in correspondingly serially        contiguous relationship in the fragmentation portion and in        generally helical relationship with respect to the longitudinal        axis.

For example, the fragmentation portion is configured for fragmentinginto said at least one set of serially adjacent fragments in generallyhelical relationship with respect to the longitudinal axis, in responseto detonation of the explosive charge.

According to the first aspect of the presently disclosed subject matter,there is also provided a fragmentation warhead configured for beingmounted in a carrier vehicle, the warhead having a longitudinal axis andcomprising:

-   -   a shell extending along said longitudinal axis and comprising a        fixed shell portion and a fragmentation portion, and defining        therebetween a cavity for accommodating therein an explosive        charge;    -   the fragmentation portion configured for fragmenting into at        least one set of serially adjacent fragments in generally        helical relationship with respect to the longitudinal axis, in        response to detonation of the explosive charge.

For example, prior to said detonation, said at least one set of seriallyadjacent fragments are correspondingly serially contiguous relationshipin said fragmentation portion and in generally helical relationship withrespect to the longitudinal axis.

Additionally or alternatively, the fragmentation portion is configuredfor fragmenting into a plurality of laterally adjacent said sets ofserially adjacent fragments in generally helical relationship withrespect to the longitudinal axis, in response to detonation of theexplosive charge.

Additionally or alternatively, the fragmentation portion is configuredfor fragmenting into three laterally adjacent said sets, each said setcomprising between 30 and 50 said serially adjacent fragments ingenerally helical relationship with respect to the longitudinal axis, inresponse to detonation of the explosive charge.

Additionally or alternatively, said fixed shell portion is configured sothat upon initiation of detonation of the explosive charge, shockwavespropagating therefrom are directed via said fixed shell portion thetowards said fragmentation portion.

Additionally or alternatively, said fixed shell portion has rotationalsymmetry about said longitudinal axis.

Additionally or alternatively, said fixed shell portion has a generallytubular configuration.

Additionally or alternatively, each said set of serially adjacentfragments in correspondingly serially contiguous relationship in saidfragmentation portion and in generally helical relationship with respectto the longitudinal axis is oriented at predetermined helix angle withrespect to said longitudinal axis. For example, said helix angle ispredetermined such that upon said detonation, the respective saidfragments of each said set are spread over an imaginary cylindricalsurface along a distance of between about 2 m to about 4 m over saidcylindrical surface, at a corresponding radial distance of between 4 mand 8 m, respectively, from said longitudinal axis, while ensuring aspacing of not greater than 0.1 m between adjacent fragments at saidradial distance. For example, said helix angle is between 2.5° and 3°.

Additionally or alternatively, said fragmentation portion is formed as aplurality of axially serially adjacent fragmentation portion sections,each said fragmentation portion section comprising a plurality of saidfragments in lateral contiguous (abutting) relationship, and whereinsaid plurality of fragments of each successive said fragmentationportion section along said longitudinal axis is angularly displacedabout the longitudinal axis with respect to the respective saidplurality of fragments of the previous said fragmentation portionsection. For example, a respective said plurality of fragments of afirst said fragmentation portion section at one longitudinal end of saidfragmentation portion is angularly displaced about the longitudinal axiswith respect to a respective said plurality of fragments of a last saidfragmentation portion section at another longitudinal end of saidfragmentation portion by an angular displacement of between 25° and 35°.For example, the respective said plurality of fragments of each pair ofsuccessive said fragmentation portion sections are angularly displacedfrom one another with reference to (along) a plane orthogonal to thelongitudinal axis by a fragmentation set dispersion angle. For examplesaid fragmentation set dispersion angle is between 0.5° and 0.7°.

Additionally or alternatively, said fragmentation portion is formed as agenerally helical band with respect to said longitudinal axis, andwherein said fixed shell portion comprises a generally helical slotcomplementary to said helical band. For example, said generally helicalband is projectable to a plane to provide a two dimensionalparallelogram pattern of said fragments. For example, said parallelogrampattern having a base corresponding to the width of three saidfragments, and a height corresponding to the axial length of thefragmentation warhead. For example, each said fragment has a weight ofbetween 25 g and 35 g.

Additionally or alternatively, said fragmentation portion contains atotal number of said fragments between 100 said fragments and 200 saidfragments.

Additionally or alternatively, said fragmentation portion contains atotal number of said fragments having an aggregate weight of between 3.5Kg and 4.5 Kg.

Additionally or alternatively, each said fragment has a plan shape inthe form of a parallelogram.

Additionally or alternatively, following detonation of the explosivecharge, said fragments of each said set are serially spaced at a spacingbetween each adjacent pair of said fragments less than a diameter of aGRAD rocket warhead at a distance less than 7.5 m from said longitudinalaxis.

Additionally or alternatively, following detonation of the explosivecharge, said fragments of each said set are serially spaced at a spacingless than 0.1 m at a distance between 4.5 m and 7.5 m from saidlongitudinal axis.

Additionally or alternatively, following detonation of the explosivecharge, each said set of fragments is spaced from an adjacent said setof fragments at a spacing less than an axial length of a GRAD rocketwarhead at a distance less than 7.5 m from said longitudinal axis.

Additionally or alternatively, following detonation of the explosivecharge, each said set of fragments is spaced from an adjacent said setof fragments at a spacing less than 0.50 m at a distance between 4.5 mand 7.5 m from said longitudinal axis.

Additionally or alternatively, each said fragment is capable ofneutralizing a flying object by impacting a kill zone thereof. Forexample, said kill zone has a length of 0.50 m and a width of 0.10 m.For example, said flying object is any one of: a rocket, a GRAD rocket,a UAV, a manned aircraft, a cruise missile.

According to the first aspect of the presently disclosed subject matter,there is also provided a carrier vehicle for a fragmentation warhead,comprising:

-   -   the fragmentation warhead as defined above for this aspect of        the presently disclosed subject matter;    -   an uplink for receiving commands from a control;    -   a proximity fuse operatively connected to the fragmentation        warhead and configured for detonating the warhead at a        predetermined spacing between the carrier vehicle and a flying        object;    -   the carrier vehicle being maneuverable at least responsive to        receiving said commands.

For example, the carrier vehicle is configured for being mounted in abooster stage.

For example, said proximity fuse is configured for generating two flatlaser beams and for fusion time determination based on reflectionsreceived from said beams.

For example, said longitudinal axis is parallel to a longitudinal axisof the carrier vehicle.

For example, said uplink comprises a receiver for receiving data orsignals relating to PIP, target and carrier vehicle state vectors,and/or relative state vectors between target and carrier vehicle.

For example, the carrier vehicle comprises a plurality of pivotablevanes for steering (maneuvering) said carrier vehicle.

For example, the carrier vehicle further comprises a homing sensor,configured for autonomously homing onto a target.

According to the first aspect of the presently disclosed subject matter,there is also provided a missile for intercepting a flying object,comprising:

-   -   (a) a carrier vehicle for a fragmentation warhead as defined        above for this aspect of the presently disclosed subject matter        and    -   (b) a booster stage for propelling the carrier vehicle along a        desired trajectory.

For example, said booster stage is based on a GRAD rocket system orwherein said booster stage comprises a GRAD rocket motor.

According to the first aspect of the presently disclosed subject matter,there is also provided an interception system comprising:

-   -   a missile battery including at least one missile for        intercepting a flying object as defined above for this aspect of        the presently disclosed subject matter;    -   a radar system for detecting and tracking at least one said        flying object;    -   a computer system for determining a predicted impact point (PIP)        for the missile;    -   a communications uplink to provide maneuvering data to the        carrier vehicle during flight thereof to intercept the        respective said flying object at the respective predicted        interception point PIP.

For example, the system is configured for causing the carrier vehicle tobe selectively oriented at a desired relative angle to a flight path ofthe flying object at the predicted interception point (PIP).

For example, said relative angle is between 10° and 12°, or wherein saidrelative angle is between −20° to +40° when in pursuit interceptionscenario, or wherein said relative angle is between +160° to +220° forhead on interception scenario.

According to the first aspect of the presently disclosed subject matter,there is also provided a method for intercepting a flying object,comprising:

-   -   (i) providing a missile as defined above for this aspect of the        presently disclosed subject matter;    -   (ii) using the booster stage to selectively launch the carrier        vehicle along an intercept trajectory with respect to the flying        object;    -   (iii) maneuvering the carrier vehicle into proximity with the        flying object;    -   (iv) detecting the flying object within a minimum range with        respect to the fragmentation warhead via the proximity fuse; and    -   (v) detonating the explosive charge responsive to step (iii).

For example, in step (iii) the carrier vehicle is oriented at a relativeangle to the flying object at a predicted interception point (PIP).

For example, said relative angle is between 10° and 12°, or wherein saidrelative angle is between −20° to +40° when in pursuit interceptionscenario, or wherein said relative angle is between +160° to +220° forhead on interception scenario.

For example, the method includes providing a spacing between the carriervehicle and the flying object of between 4.5 m and 7.5 m at the PIP,and/or wherein the fragmentation portion is facing the flying object atthe PIP

According to the first aspect of the presently disclosed subject matter,there is also provided an interception missile comprising afragmentation warhead and configured for being maneuvered to a predictedinterception point (PIP) for intercepting within a probability envelopea flying object having a kill zone of known dimensions, thefragmentation warhead being configured for selectively providing aplurality of fragments directed towards said probability envelope suchthat the spacing between any two adjacent said fragments within theprobability envelope is less than at least one of said known dimensionsto ensure that at least one said fragment impacts said kill zone withinsaid probability envelope, wherein each said fragment is capable ofneutralizing the flying object by impacting said kill zone.

For example, said kill zone has a length of 0.50 m and a width of 0.10m.

For example, said flying object is any one of: a rocket, a GRAD rocket;a UAV, a manned air vehicle, cruise missile.

According to the first aspect of the presently disclosed subject matter,there is also provided a fragmentation warhead configured for beingmounted in a carrier vehicle, the warhead having a longitudinaldimension along a longitudinal axis and configured for selectivelyproviding a plurality of fragments directed towards a target area spacedat an interception spacing from said warhead to provide a fragmentationpattern on the target area including at least one set of said fragmentsin a spaced linear relationship extending to a length dimension greaterthan said longitudinal dimension, wherein adjacent said fragments ineach said set at the target area are spaced at a respectivefragmentation spacing that is within the dimensions of a kill zone of anintended target, wherein each said fragment is capable of neutralizingthe intended target by impacting said kill zone.

For example, said kill zone has a length of 0.50 m and a width of 0.10m.

For example, said flying object is any one of: a rocket, a GRAD rocket;a UAV, a manned air vehicle, cruise missile.

BRIEF DESCRIPTION OF THE DRAWINGS

In order to better understand the subject matter that is disclosedherein and to exemplify how it may be carried out in practice, severalexamples will now be described, by way of non-limiting examples only,with reference to the accompanying drawings, in which:

FIG. 1 is a schematic representation illustrating in side view aninterception missile according to an example of the presently disclosedsubject matter.

FIG. 2 shows in partially fragmented view the carrier vehicle of theinterception missile of the example of FIG. 1.

FIGS. 3A, 3B illustrate schematically the aft portion of the carriervehicle of the example of FIG. 1; FIG. 3C illustrates airborne mainsection system architecture, in accordance with certain examples.

FIG. 4 shows in isometric view a fragmentation warhead of the example ofFIG. 1

FIG. 5 shows in cross-sectional view the fragmentation warhead of theexample of FIG. 4;

FIG. 5A shows in cross-sectional view a variation of the fragmentationwarhead example of FIG. 5.

FIG. 6 schematically illustrates the fragmentation pattern of one set offragments of the fragmentation warhead of the example of FIG. 4; FIG. 6Aschematically illustrates the fragmentation patterns three sets offragments of the fragmentation warhead of the example of FIG. 4.

FIG. 7 illustrates the projection onto a plane of the fragmentationportion of the example of FIG. 4.

FIG. 8 illustrates an example or a target incoming rocket warhead.

FIG. 9 schematically illustrates operation of the proximity fuse of theexample of FIG. 1; FIG. 9A schematically illustrates degree ofuncertainty associated with operation of the proximity fuse.

FIG. 10 schematically illustrates divergence of fragments of the exampleof FIG. 5.

FIG. 11 schematically illustrates a probability envelope according to anex ample thereof.

FIG. 12A illustrates a generalized detection and tracking scenario thatis devoid of utilization of a seeker head, in accordance with certainexamples of the presently disclosed subject matter;

FIG. 12B illustrates ground control system architecture, in accordancewith certain examples of the presently disclosed subject matter;

FIG. 13 illustrates schematically a block diagram for implementingaccurate synchronization between clocks among sensors, in accordancewith certain examples of the presently disclosed subject matter;

FIG. 14 is a chart illustrating synchronization results, in accordancewith certain examples of the presently disclosed subject matter;

FIG. 15 illustrates a typical interception scenario, in accordance withcertain examples of the presently disclosed subject matter;

FIG. 16 illustrates an exemplary sensor array layout in an interceptionscenario, in accordance with certain examples of the presently disclosedsubject matter;

FIG. 17 illustrates a deployment scheme of a combined (radar andpassive) sensor synchronized at a high accuracy, in accordance withcertain examples of the presently disclosed subject matter;

FIGS. 18, 18A, 18B illustrate measurement of target at Y coordinate;

FIG. 19 illustrates measurement of target at X coordinate; and

FIG. 20 illustrates a sequence of operations, in accordance with certainexamples of the presently disclosed subject matter.

DETAILED DESCRIPTION OF THE FIGURES

Referring to FIG. 1, an interception missile (for example, forintercepting a flying object) according to a first example of thepresently disclosed subject matter is generally designated withreference numeral 100, and comprises a carrier vehicle 102 for afragmentation warhead, and a booster stage 101.

The booster stage 101 is configured for launching and propelling thecarrier vehicle 102 along a desired trajectory, and comprises a suitablepropulsion system (not shown). While at least in this example thepropulsion system comprises a solid fuel rocket motor and solidpropellant accommodated within the external casing 109, in alternativevariations of this example the propulsion system can instead oradditionally include one or more liquid fuel rocket engines, forexample. The booster stage 101 further comprises a plurality of fins 103for stabilizing the missile 100, for example during the boost and/orcoast phases of the flight, and until the carrier vehicle 102 separatesfrom the booster stage 101. In this example, the fins 103 are configuredas wrap-around fins, being deployable from a stowed configuration to adeployed configuration. In the stowed configuration the fins are wrappedaround the periphery of the aft section 111 of the booster stage 101,providing a compact geometry, which can be useful for storage and/or forlaunching the missile 100 from a launch tube of internal diameterslightly greater than the maximum outer diameter of the missile 100. Inthe deployed configuration (as illustrated in FIG. 1) the fins 103 areprojecting outwardly, generally radially with respect to thelongitudinal axis AB of the booster stage 101. In alternative variationsof this example, different stabilizing systems can be used, for examplefixed fins.

The forward end 108 of the booster stage 101 is configured for enablingthe carrier vehicle 102 to be releasably mounted thereto, for examplevia explosive bolts.

In operation, the booster stage 101 serves to boost the missile 100 toacquire maximum kinetic energy to propel the missile along a desiredtrajectory, and to increase the ballistic coefficient of the missile 100during the coasting phase. Stage separation then occurs between thecarrier vehicle 102 and the booster stage 101, in which the boosterstage 101 is discarded, and the carrier vehicle continues along a guidedtrajectory to the predicted interception point (PIP). In at least thisexample the missile 100 is not initially guided, but rather is launchedwith an initial azimuth and elevation that is calculated to bring thecarrier vehicle into proximity with the PIP, and guidance is providedonly after stage separation.

In this example, the booster stage 101 is based on a GRAD type rocket,which is already developed and mass-produced. For example, the boosterstage 101 can include at least a part of such a GRAD type rocket,suitably modified if necessary to allow for the mounting of, andsubsequently to the selective separation of, the carrier vehicle 102. Insuch examples, the unit cost of each booster stage 101 is comparativelylow, as compared with the unit cost of a booster stage that is developedspecifically for use with the carrier vehicle 102. As will becomeclearer below, this feature of utilizing a GRAD type rocket forproviding the booster stage 101 contributes to minimizing the economicunit cost of each missile 100, so that in at least some examples, such aunit cost can be comparable to, i.e., at least within the same order ofmagnitude as, the unit costs of GRAD type rockets that are intended tobe intercepted and neutralized by the missile 100. Furthermore, in atleast some examples the missile 100 can be launched using launchersdesigned for regular GRAD type rockets, thereby reducing costs ofoperation even further.

By neutralizing an object is meant herein to refer to destroying theobject, or to otherwise render the object ineffective, or to at leastsignificantly reduce the lethality of the object.

GRAD type rockets are well known in the art and can be defined as a MLRS(multiple launch rocket system) class of rockets that are mass producedas rockets that have a relative low economic unit cost.

An example of such a GRAD type rocket is the M-21OF rocket, produced inRussia, and typically launched from a BM-21 launch vehicle that includesa multiple rocket launcher. Some examples of such a GRAD type missilehave an external diameter of 122 mm, length 2.8 m, take-off weight ofabout 70 Kg including a payload (warhead) weight of about 19 Kg; somevariants have a ground-to-ground range of about 12 km, while others havea 20 km range. Improved variants of GRAD rockets have range of about 40km

Referring also to FIG. 2, the carrier vehicle 102 comprises a warheadmodule 104, proximity fuse 105, and aft module 106.

The proximity fuse 105 in this example comprises a laser detector, whichgenerates two laser beams, angularly displaced from one another withrespect to the axis AC, and a laser detector for detecting reflectedlaser beams. Each laser beam is formed as a flat beam. The proximityfuse 105 is configured for fusion time determination based onreflections received from the laser beams. The proximity fuse 105operates (directly or via a mission computer) to detonate the explosivecharge 360 at a calculated time interval after a flying object crossesboth beams, as will become clearer below. In any case proximity fusesare well known in the art, and include, for example proximity sensorsprovided by L3 (USA), or any other suitable range finder systems basedon LADAR or radar techniques. In alternative variations of this example,different types of proximity fuses may be used, mutatis mutandis.

Referring also to FIG. 3A, the aft module 106 comprises a battery 201,steering system 202, navigation system 203, and communication system204.

The steering system 202 includes drivers (not shown), actuators 205, andsteering fins 107. In this example, the fins 107 are pivotably mountedto the aft module 106, being deployable from a stowed configuration to adeployed configuration. In the stowed configuration (see for exampleFIG. 1) the fins 107 are accommodated in axial slots 107A (FIG. 2)provided in the aft module 106. In the deployed configuration (asillustrated in FIGS. 2, 3A, 3B), the fins 107 are pivoted out of theirrespective slots to project radially with respect to the longitudinalaxis AC of the carrier vehicle 102. Once deployed, the fins can bepivoted about respective axes VA that radially project from longitudinalaxis AC of the carrier vehicle 102 to provide one or more of pitchcontrol moments, roll control moments, and yaw control moments withrespect to the longitudinal axis AC of the carrier vehicle 102. Inalternative variations of this example, different steering systems canbe used, for example foldable pivotable fins (e.g. wrap-around pivotablefins) or permanently deployed pivotable fins.

The navigation system 203 includes airborne computer 206 and an inertialunit (not shown) configured for measuring angular displacement todetermine the spatial orientation of the missile 100, and in particularof the carrier vehicle 102 with respect to three orthogonal axes (forexample the pitch roll and yaw axes of the carrier vehicle 102).Optionally, the navigation system 203, in particular the inertial unit,can comprise accelerometers along each of the three orthogonal axes forincreasing accuracy of the determination of the missile spatiallocation, in particular for increasing accuracy of the determination ofthe spatial location (and spatial orientation) of the carrier vehicle102, with reference to the three orthogonal axes, between sequentialuplink communications.

The airborne computer 206 is configured for performing a number of tasksincluding at least the following:

-   -   guidance task for guiding the carrier vehicle 102 to a predicted        interception point (PIP);    -   navigation task for determining the spatial position and spatial        orientation of the carrier vehicle 102 during flight;    -   control task for close loop control of the steering mechanism,        including providing a desired spatial orientation of the carrier        vehicle 102 at the PIP;    -   activation of stage separation between the carrier vehicle 102        and the booster stage 101;    -   activation of the warhead 104 including the proximity fuse 105.

The communication system 204 includes uplink receiver (not shown) andantenna 207, for providing updated control commands or information tothe missile 100, and in particular to the carrier vehicle 102, duringflight.

Referring to FIG. 3C, an airborne main system architecture isillustrated in accordance with at least some examples of the presentlydisclosed subject matter, in which the steering system 202 includesdrivers 2001 configured to command actuators 2002 for actuating steeringfins 2003 (for example fins 107), as known per se. Airborne computer2004 (forming part of the navigation system) is powered by battery 2005and receives feedback indicative of the steering fins 2003 (by means ofthe steering drivers 2001) and the body angle rate measurements module2006 (forming part of navigation system) determining an angular positionof the carrier vehicle 102 in three mutually orthogonal axes. Theairborne system further receives inputs through the uplink receiver 2007and associated antenna (forming part of the communication system frommain external source (for example ground sensors) for updating thecarrier vehicle 102 during flight. The airborne computer 2004 processesall these data for:

-   -   (i) guidance task for diverting or otherwise guiding the carrier        vehicle 102 towards the PIP by generating steering commands that        are sent to the steering fins 2003 through the steering drivers;    -   (ii) navigation tasks for determining the spatial position and        spatial orientation of the carrier vehicle 102 during the flight        (based on the angle rate measurement module 2006);    -   (iii) control task for close loop control of the steering        mechanism, including providing a desired spatial orientation of        the carrier vehicle 102 at the PIP;    -   (iv) activation of stage separation between the carrier vehicle        102 and the booster stage 101;    -   (v) activation of the warhead including the proximity fuse 2008        (for example fuse 105).

It is to be noted that the presently disclosed subject matter is notbound by the specified system architecture.

In operation, the airborne computer 2004 receives updated navigationdata from the on board inertial measurement unit and from an externalmeasurement of the location of the carrier vehicle by the uplinkcommunication channel (using the uplink receiver 2007 of communicationsystem).

In accordance with at least certain examples, the guidance rules whichcontrol the steering of the carrier vehicle 102 towards the PIP cancomply (but not necessarily) for example with a known per seproportional navigation paradigm.

Referring to FIGS. 2 and 4, the warhead module 104 accommodatesfragmentation warhead 300.

Thus, fragmentation warhead 300 is configured for being mounted in thecarrier vehicle 102. The fragmentation warhead has a longitudinal axisLA, which in this example is co-axial with longitudinal axis LC of thecarrier vehicle 102, and also co-axial with the longitudinal axis LB ofthe booster 101 (while the carrier vehicle 102 is mounted to the booster101). However, in alternative variations of this example, thelongitudinal axis LA, can be parallel with respect to, or alternativelyat a non-zero angle with respect to, longitudinal axis LC of the carriervehicle 102 and/or with respect to the longitudinal axis LB of thebooster 101 (while the carrier vehicle 102 is mounted to the booster101).

Referring in particular to FIG. 4 and FIG. 5, the fragmentation warhead300 includes a shell 320 extending along said longitudinal axis LA andcomprising a fixed shell portion 310 and a fragmentation portion 350. Acavity 330 is defined between the fixed shell portion 310 and afragmentation portion 350, in which there is accommodated an explosivecharge 360.

The fixed shell portion 310 is configured so that upon initiation ofdetonation of the explosive charge 360, shockwaves propagating therefromare directed via the fixed shell portion 310 the towards saidfragmentation portion 350. Thus, on detonation of the explosive charge,the fixed shell portion 310 retains its mechanical integrity and directsthe blast of the explosion towards the fragmentation portion 350,causing the fragmentation portion to fragment into the individualfragments 355, and causing the fragments to be ejected at high velocityin directions F1, F and F2.

As will become clearer herein, and referring to FIGS. 6 and 6A, thefragmentation portion 350 is configured for fragmenting into threelaterally adjacent sets 370 of individual fragments 355, wherein eachset 370 comprises a plurality of serially adjacent but spaced fragments355 in generally helical relationship with respect to the longitudinalaxis LA, in response to detonation of the explosive charge 360—for thesake of clarity FIG. 6 shows only one such set 370, while FIG. 6A showsthe three sets 370. In alternative variations of this example, thefragmentation portion 350 is configured for fragmenting into only oneset 370 of fragments 355, or for fragmenting into two laterally adjacentsets 370 of fragments 355, or for fragmenting into a plurality(including more than three) laterally adjacent sets 370 of fragments355, wherein each set 370 comprises a plurality of serially adjacentfragments 355 in generally helical relationship with respect to thelongitudinal axis LA, in response to detonation of the explosive charge360.

The shell 320, and in particular the fixed shell portion 310 hasrotational symmetry about the longitudinal axis LA. In this example, theshell 320, and in particular the fixed shell portion 310 is generallytubular, with end walls 321, 322, which in this example are generallyflat in the vicinity of the fragmentation portion 350. However, inalternative variations of this example, the shell 320, and in particularthe fixed shell portion 310 can have a different rotational symmetry,for example in the form of a bulging cylinder or capped sphere, and/orthe end walls can be non-flat, and/or the fixed shell portion can have apolygonal cross-section, for example.

The fragmentation portion 350 is formed as a generally helical band 390with respect to the longitudinal axis LA, and the fixed shell portion310 comprises a generally helical slot defined by edges 311, 312 (FIG.5) complementary to the helical band 390. The helical band 390 in thisexample comprises three laterally adjacent sets 370 of fragments 355(separately referred to as sets 370A, 370B, 370C), wherein each set 370comprises a plurality of serially adjacent fragments 355 incorrespondingly serially contiguous relationship in the fragmentationportion 350 and in generally helical relationship with respect to thelongitudinal axis LA.

Thus, in each set 370, the fragments 355 are contiguous prior todetonation of the explosive charge 360, and are subsequently spacedapart after detonation and ejection of the fragments 355 from thefragmentation warhead 300.

As illustrated in FIG. 7, the generally helical band 390 is projectableto a plane to provide a flat projection of the band, designated 390′ inthis figure. The projected band 390′ forms a two dimensional pattern ofthe contiguous fragments 355, the pattern being in the form of aparallelogram having a base B and height H. Correspondingly, in thisexample, each fragment 355 also has a parallelogram planform.

Each set 370 is thus oriented at a helix angle Φ with respect to thelongitudinal axis LA. In this example, the helix angle Φ is such thatthe respective fragments 355 of each set 370, after detonation of theexplosive charge 360 and ejection of the fragments 355 from thefragmentation warhead, are spread over an imaginary cylindrical surfaceCS (see FIGS. 6, 6A) along a distance on this surface CS of betweenabout 2 m to about 4 m, at a corresponding radial distance of between4.5 m and 7.5 m, respectively, from said longitudinal axis, whileensuring a spacing of not greater than 0.10 m between adjacent fragments355 at this radial distance. Thus, each set 370 requires between 20 and40 fragments.

For example, the helix angle Φ is between 2.5° and 3°.

In this example, the base B corresponds to the width w of threefragments 355, and a height H corresponding to the axial length L of thefragmentation warhead 300. In this example, the width w of each fragment355 is about 24 mm, and thus the base B is about 72 mm. The thickness ofeach fragment 355 can be about 5 mm, for example.

The fragmentation portion 350 is formed as a plurality of axiallyadjacent fragmentation portion sections 375, each fragmentation portionsection 375 comprising a number of fragments 355 in lateral adjacent,abutting or contiguous relationship, In this example, each fragmentationportion section 375 has three fragments 355: a first fragment 355belonging to the first set 370A, a second fragment 355 belonging to thesecond set 370B, and a third segment 355 belonging to the third set370C. In this example, there are 38 fragmentation portion sections 375along the axis LA, corresponding to the 38 fragments of each set 370.Each fragment has a length l of about 17 mm, and thus the length S ofthe long size of the projected band 390′ is about 646 mm. In alternativevariations of this example, each set can have more than or less than 38fragments 355.

In alternative variations of this example, the fragmentation portion cancomprise three sets 370 of 40 fragments each, and/or the fragments 355can be rhombic-shaped with width w 17.6 mm and length 17.6 mm andthickness 6.6 mm; thus the corresponding base B is 53 mm.

The plurality of fragments 355 of each successive fragmentation portionsection 375 along the longitudinal axis LA is angularly displaced aboutthe longitudinal axis LA with respect to the respective plurality offragments 355 of the previous fragmentation portion section 375, therebyproviding the generally helical relationship of the fragments 355 ineach set 370 with respect to the longitudinal axis LA.

In this example, and referring to FIG. 4, the plurality of fragments 355of a first fragmentation portion section 375 at one longitudinal end ofthe fragmentation portion 350 (corresponding to end wall 321) isangularly displaced about the longitudinal axis LA with respect to theplurality of fragments 355 of a last fragmentation portion section 375at another longitudinal end of the fragmentation portion 350corresponding to end wall 322) by an angular displacement correspondingto an aggregate dispersion angle θ. In this example, angle θ is between25° and 35°, preferably about 30°, though in alternative variation ofthis example, angle θ can instead be different, for example between 18°and 22°. The aggregate dispersion angle θ determines the density offragments 355 at predetermined distances from the axis LA. For example,the larger the aggregate dispersion angle θ, the larger the spread ofthe fragments 355 at a given radial distance, but the greater thespacing between adjacent fragments 355, for the same number offragments. The choice of aggregate dispersion angle θ depends on theaccuracy of determination of relative distance and position between theflying object and the carrier vehicle, and the size of killing zone ofthe flying object.

Thus, in this example, the respective plurality of fragments 355 of eachpair of successive fragmentation portion sections 375 are angularlydisplaced from one another with reference to (i.e., along) a planeorthogonal to the longitudinal axis LA by a fragmentation set dispersionangle θ′. Such fragmentation set dispersion angle θ′ is between about0.5° and about 1°. For example, fragmentation set dispersion angle θ′ isbetween about 0.75° and about 0.8° for an aggregate dispersion angle θof 30°, for achieving a maximum spacing of 0.1 m between two adjacentfragments 355 at a radial distance of about 7.5 m, and between about0.45° and about 0.55° for an aggregate dispersion angle θ of 20° forachieving a maximum spacing of 0.1 m between two adjacent fragments 355at a radial distance of about 10 m, assuming 40 fragments per set 370.

In this example, and referring again to FIG. 5, the respective fragments355 of each fragmentation portion section 375 can be coplanar, or ingeneral convex relationship with respect to cavity 330 (i.e., projectingconvexly away from cavity 330), or in general concave relationship withrespect to cavity 330 (i.e., projecting convexly toward cavity 330),depending on the magnitude of the angular dispersion desired between thefragment sets 370. Thus, while the central fragment 355 is facingdirection F, each of the two outer fragments 355 are facing indirections F1 and F2, respectively, angularly displaced from direction Fby angle α. For relatively small angular dispersion α, of less than 5°for example, the respective fragments 355 of each fragmentation portionsection 375 are in general concave relationship with respect to cavity330 (i.e., projecting convexly toward cavity 330), as illustrated inFIG. 5A. without being bound to theory, it is considered that the actualdirection of ejection (angle α) of the fragments 355 of the outer sets370A, 370C is determined by an interaction of direct shockwaves with thefragments 355 originating from the explosive material 360, theshockwaves reflected by the fixed shell portion 310 to the fragments355, and boundary effects between the outer fragments 355 and the shellat edges 311, 312, and boundary effects between adjacent fragments 355.Thus, for example, in variations of this example where the adjacentfragments 355 of the three sets 370 are coplanar, the fragments can beejected with a dispersion angle α greater than about 5°.

The curvature formed by the respective fragments 355 of eachfragmentation portion section 375 is less than the curvature of thefixed shell portion 310, and provides a space 315 between thefragmentation portion 350 and the outer casing 316 of the warhead module104.

In this example, each fragment 355 has a weight of between 25 g and 35g, for example about 30 g. However, in alternative variations of thisexample, each fragment can instead have a weight of between 2 g and 50g, depending on the nature of the target, and the specific configurationof the respective carrier vehicle and fragmentation warhead.

In this example, the fragmentation warhead 300, in particular thefragmentation portion 355 contains a total number of said fragmentshaving an aggregate weight of between 3 Kg and 4.5 Kg, for example 3.42Kg. However, in alternative variations of this example, the fragments355 can have an aggregate weight of between 2.4 Kg and 5 Kg, for example

In general, the weight of the explosive material 360 at least matchesthat of all the fragments 355, say another 4 Kg, and the weight of theshell fixed portion 310 also generally matches that of the fragments.Thus, for an aggregate weight of about 4 Kg, the fragmentation warheadis about 12 Kg. in addition, the remainder of the carrier vehicle 102(structure, avionics etc) can weigh another 12 Kg.

Thus, the net weight of the fragmentation portion 350 allows the carriervehicle 102 to weigh 24 Kg or less, and allows this to be carried by abooster stage 101 that is based on a Grad type rocket, in which theconventional Grad warhead has a comparable weight of about 19 Kg.

In this example, the fragmentation warhead 300, in particular thefragmentation portion 350 contains a total number of fragments 355 ofbetween 100 fragments and 200 fragments, for example 114 fragments.However, in alternative variations of this example, the fragmentationwarhead 300, in particular the fragmentation portion 350 can contain atotal number of fragments 355 of between 1,200 light fragments (of 2 greach) and 100 fragments (of 50 gr each) while maintaining the sameaggregate weight between 2.5 Kg and 5 Kg for the fragmentation portion350.

In this example, each fragment 355 is configured to be capable toneutralize the warhead of an incoming grad-type rocket, under thefollowing conditions:

-   -   (A) the respective fragment 355 is ejected from the        fragmentation warhead 300 via detonation of the explosive charge        360 to provide an impact velocity with respect to the warhead of        an incoming grad-type rocket sufficient to impact and penetrate        the warhead of an incoming grad-type rocket;    -   (B) the respective fragment 355 impacts the warhead of an        incoming grad-type rocket within a predefined kill zone defined        with respect to the warhead of an incoming grad-type rocket,        such that penetration of the fragment 355 in this kill zone        neutralizes the warhead.

In other words, each fragment 355 is capable of neutralizing theincoming warhead by impacting the kill zone.

Regarding condition (A), this can be achieved by providing therespective fragment 355 with the required momentum. For example this canbe achieved by detonating the explosive charge 360 when the warhead ofan incoming grad-type rocket is within a predetermined range, and byproviding a suitable material for the explosive charge, which are wellknown in the art. For example, such a suitable impact velocity can beprovided by intercepting the warhead of an incoming grad-type rocketwithin a 10 m range, for example within a range of 4.5 m to 7.5 m, fromthe fragmentation warhead 300, and by ejecting the fragments 355 at avelocity of about 1.7 km/s. The impact velocity is determined from theejection velocity and the closing velocity between the carrier vehicleand the flying object. For example, such an impact velocity can be about2 km/s.

Regarding condition (B), and referring to FIG. 8, the warhead of anincoming grad-type rocket is designated with reference numeral 1, andtypically has a generally ogive or conical nose 2 and cylindrical aftsection 3, having an overall length dimension GL of about 700 mm, anddiameter GD of about 122 mm. The kill zone, designated KZ in thisfigure, is defined within the volume of the incoming threat, for examplein the form of warhead 1, for example as cylindrical surface or volumeof diameter KD and length KL (being smaller than the diameter GD andlength dimension GL, respectively), coaxial with the incoming warhead 1.Thus, the kill zone KZ appears as a rectangular zone when the incomingwarhead 1 is seen in side view, as in FIG. 8. For a typical Grad typerocket, the kill zone length KL can be about 500 mm, and the kill zonediameter KD can be 100 mm. This value for kill zone diameter KDcorresponds to providing a strike angle with respect to the surface ofthe incoming warhead 1 of 30° or more.

According to an aspect of the presently disclosed subject matter thefragmentation warhead 300 is configured for being maneuvered to apredicted interception point (PIP) via the carrier vehicle 102 forintercepting within a probability envelope PE a flying object having akill zone of known dimensions. For example the flying object is warhead1 with kill zone KZ of size 500 mm by 100 mm. In other words, and aswill become clearer herein, the probability envelope PE can beconsidered to be a volume in space associated with the PIP where it isdetermined, with a high degree of certainty, that the kill zone KZ is tobe found. As will become clearer herein, the fragmentation warhead 300is configured for selectively providing a plurality of fragments 355directed towards this probability envelope PE such that the spacingbetween any two adjacent fragments 355 within the probability envelopePE is less than at least one of dimensions of the kill zone KZ, toensure that at least one fragment 355 impacts the kill zone KZ withinthe probability envelope PE, thereby neutralizing the incoming warhead.

Referring to FIG. 9, it is anticipated that, at least in some examples,the carrier vehicle 102 approaches the PIP with longitudinal axis ACnon-parallel to the longitudinal axis of the warhead 1, which istypically aligned with flight path FP of the incoming warhead 1, forexample more or less aligned with the flight path FP. Alternatively, forexample, the longitudinal axis AC is angled to the longitudinal axis ofthe warhead 1 (which is typically aligned with flight path FP) at arelative angle λ up to −20° to +40° to the flight path FP for pursuitscenario, or between +160° to 220° for head-on scenario. Further, theproximity fuse 105 detects the presence of the warhead 1 in proximity tothe carrier vehicle 102, and in response thereto the explosive charge360 is detonated.

Referring to FIG. 6, it is readily appreciated that the helix angle ofthe fragments 355 of each set 370 when the fragments 355 are still onthe fragmentation warhead 350 (i.e., prior to detonation) increasesrapidly with radius from the axis LA as the fragments 355 are ejectedafter detonation. Thus, starting with a small helix angle of about 2° or3° prior to detonation, the helix angle can increase to about 80° at adistance of between 4.5 m and 7.5 m from the axis LA. Thus, the relativeangle between the axis LA and the flight path FP of about 10° (or 190°)ensures that the disposition of the ejected fragments 355 of each set370 with respect to the flight path FP ate the PIP will be approximatelyorthogonal. Referring to FIG. 11, this is the optimal angularrelationship between the fragments and the longitudinal axis of theflying object (assumed to be parallel to the flight path FP thereof).

Referring to FIG. 11, it is to be noted that increasing the helix angleresults in a decrease in the spacing d between adjacent sets 370 ofejected fragments 355, in a direction SP1 orthogonal to the direction ofspread SP2 of the fragments along each set 370. This effect can becompensated by increasing the initial dispersion angle α betweenadjacent sets of fragments 370A, 370B, 370C.

Another effect of increasing the helix angle is related to the partialoverlap between the adjacent sets along the direction of the spread SP2of the fragments along each set 370. As can be seen from FIG. 11, thespread of the ejected fragments 355 when projected onto a plane is inthe form of a parallelogram, and thus:

-   -   a central portion POR1 of the ejected fragments 355 contains        fragments from all three sets 370A, 370B, 370C in overlapping        relationship;    -   a mid-portion POR2 on each side of central portion POR1 of the        ejected fragments 355 contains fragments from two sets (370A,        370B) or (370B, 370C) in overlapping relationship;    -   an outer portion POR3 on the outer side of each mid-portion POR2        contains fragments from one set (370A or 370C)

It is possible to change the degree of overlapping between the adjacentsets 370A, 370B, 370C of ejected fragments 355 by modifying the geometryof the fragmentation warhead 300, for example by modifying the generalcylindrical geometry of warhead 300 to the form of a barrel, in whichthe diameter at the center (axially) is larger than at the longitudinalends thereof.

There is a tradeoff between:

-   -   choosing the initial dispersion angle α;    -   limiting the angular deviation between the flight path of the        carrier vehicle and the flight path of the threat, with respect        to the optimal angular deviation that ensures orthogonality        between the target body (along the longitudinal axis thereof)        and the direction of spread SP2 of the ejected fragments 355 of        each et 370;    -   the requirement for accuracy of the proximity fuse.

In at least one example, final parameters of the fragmentation warheadto provide a desired spread of fragments to provide a desired spacing d,a desired overlap, and a desired tradeoff, can be finalized followingsystem performance simulations, for example.

However, it is possible to deviate from orthogonality by up to ±30° andstill provide a lethal spread of fragments in each set 370.Non-orthogonality increases the effective spacing between the ejectedadjacent sets 370, as seen by the flying object. For example, adeviation of up to ±30° effectively increases spacing between ejectedadjacent fragments 355 of the same set 370 by up to 15%.

Thus, in a pursuit scenario, in which the carrier vehicle 102 istravelling in a similar direction to the flying object, the relativeangle can be between −20° (i.e., −30° plus 10°) to +40° (i.e., +30° plus10°) to the flight path FP. On the other hand, for head-on scenario inwhich the carrier vehicle and the flying object are travelling inopposite directions towards one another, the relative angle can bebetween +160° (i.e., −30° plus 10° plus 180°) to 220° (i.e., +30° plus10° plus 180°).

In this example, the proximity fuse 105 radiates two flat laser beams105A and 105B, angularly displaced from one another with respect to theaxis AC. As illustrated in FIG. 9, a flying object such as warhead 1,for example, in the head-on interception scenario, or the wrap aroundfins, for example, in the pursuit interception scenario, will firstcross the first beam 105A, and then the second beam 105B, and theproximity fuse registers at least the times when:

-   -   the threat (e.g., the warhead 1 in the head-on interception        scenario, or the wrap around fins in the pursuit interception        scenario) interrupts the first beam 105A;    -   the threat (e.g., the warhead 1 in the head-on interception        scenario, or the wrap around fins in the pursuit interception        scenario) interrupts the second beam 105B.

In some cases, the proximity fuse 105 also registers the time when thethreat finishes interrupting the first beam 105A, and has thus fullycleared this beam, and/or the time when the threat finishes interruptingthe second beam 105B, and has thus fully cleared this beam, and theseadditional measurements can improve the accuracy of the determination offusion time by the proximity fuse.

These times allow the proximity fuse 105 to determine when the warhead 1will be at the required distance D to the carrier vehicle 102, and thusto the fragmentation warhead 300, for example using a suitable algorithm(for example based on proportional navigation of central fragments). Atthe distance, the warhead 300 will be activated. Such an algorithmincludes other inputs, for example regarding the closing velocitiesbetween the carrier vehicle 102 and the incoming threat, as provided byan external source for example, via uplinks from communication system204. Such an algorithm can also include other inputs, for exampleaverage ejection velocity of the fragments 355, and/or the time delaybetween sending of the commands for detonation of the explosive charge360 and the actual ejection of the fragments 355, and/or estimated timedelay due to atmospheric drag of the fragments 355, to achieve arequired radial distance. The proximity fuse then operates to provide asuitable detonation signal to detonate the explosive charge 360 atclosing distance D, essentially at a calculated time interval after thethreat crosses the two laser beams.

The guidance system also maneuvers the carrier vehicle 102, for exampleby providing suitable roll moments, so that the fragmentation portion350 is facing towards the flight path FP, and this is carried out priorto operation of the proximity fuse 105.

Nevertheless, there can be a degree of uncertainty as to the location ofthe warhead 1 in first direction along the flight path FP, and this isreferred to herein as the first degree of uncertainty DEG1 asillustrated in FIG. 9(a). A source for this first degree of uncertaintyDEG1 is that it is unknown whether the laser beams are impinging thefront end or the aft end of the threat, and thus it is prudent toprovide fragments 355 not just at the point P on the flight path FPcorresponding to the closing distance D, but also at a point P1 ahead ofP, and also at a point P2 aft of P. Forward point P1 is provided to takeaccount of the possibility that the aft part of the warhead 1 is atpoint P and thus most of the warhead 1 (including the kill zone KZ) isforward thereof. Aft point P2 is provided to take account of thepossibility that the nose of the warhead 1 is at point P and thus mostof the warhead 1 (including the kill zone KZ) is aft thereof. Spacingsd, between point P2 and P, and between P and P1, are each set to be notgreater than the axial length of the kill zone along the flight path FPdirection, i.e., not greater than lengthy KL, in this example notgreater than 0.5 m. This axial spread of 3 fragments 355 along theflight path FP essentially ensures hitting the kill zone KZ with respectto the axial direction (flight path FP) of the warhead 1, wherein thekill zone KZ can be in an axial range k of (2*d+2*KL), which is ±(d+KL)of point P (axially), in this example ±0.75 m to ±1 m of point P(axially). This axial range k thus defines a first dimension of theprobability envelope PE along the direction of the flight path FP.

In this example, it is further anticipated that the carrier vehicle 102will be maneuvered to the PIP, such that the closing distance D isbetween 4.5 m and about 7.5 m. Referring again to FIG. 5, the centralfragment 355 is facing direction F, and each of the two outer fragments355 are facing in directions F1 and F2, respectively, angularlydisplaced from direction F by angle α. Referring also to FIG. 10, angleα can be chosen such that the lateral spread s of each outer fragment355 with respect to the central fragment 355 does not exceed spacing d,i.e. is not greater than the length KL, which in this example is 0.5 m.Thus, with the maximum spread s set at 0.5 m at a maximum closingdistance D, angle α is arctan ((0.5/sin (helix angle))/7.5), which isminimum about 3.8°, the helix angle being that of the ejected set 370 offragments 355 at this radius. At the same time, this value for angle αprovides a lateral spread s of about 0.3 m at the minimum closingdistance D of 4.5 m.

There can be another degree of uncertainty as to the location of thewarhead 1 in a second direction orthogonal to the flight path FP, andthis is referred to herein as the second degree of uncertainty DEG2. Asource for this second degree of uncertainty DEG2 is that the lasers ofthe proximity fuse 105 can have an angular spread about the axis AC (seeFIG. 2) of the carrier vehicle 102, and it is unknown whether a forwardpart or an aft part of each laser beam is impinging the threat, and thusit is prudent to provide a plurality of fragments 355 at each of thepoints P1, P and P2, wherein each plurality of fragments is spread inthe second direction (generally orthogonal to the axial or flight pathFP direction) to cover for this uncertainty. At the same time, thespacing t between fragments 355 in each plurality of fragments 355 mustbe such that at least one such fragment will hit a kill zone KZ foundtherein. Thus, this spacing t is not greater than the lateral dimensionof the kill zone, i.e., not greater than kill zone diameter KD at themaximum closing distance D. In this example, the kill zone diameter KDis 100 mm, and this defines the maximum spacing t at maximum closingdistance D of 7.5 m. Thus, if each set 370 has a plurality of mfragments 355, the full lateral spread f of the fragments 355 at each ofthe points P1, P and P2, is thus m*t.

This lateral spread f thus defines a second dimension of the probabilityenvelope PE generally orthogonal to the direction of the flight path FP.

As illustrated in FIGS. 6 and 6A, the general helical configuration ofeach set 370 of fragments 355 in the fragmentation portion 350 allows arelatively large lateral spread f to be obtained from a fragmentationwarhead of much smaller axial length L along axis LA. For a given axiallength L of the fragmentation warhead 300 and for a given fragment size,the larger the aggregate dispersion angle θ, the greater the lateralspread f and the greater the spacing t between fragments 355.

At the same time, and referring also to FIGS. 4 and 5, the generalhelical configuration of each set 370 with the respective fragments 355in the outer sets 370A, 370C facing directions at an angle α to therespective fragments of the central set 370B provides a controlledspacing between the adjacent sets 370 of fragments 355 corresponding tospacings d. The greater the angle α, the larger the spacing s betweenadjacent sets 370 of fragments for a given closing distance D.

The range of closing distance D between the minimum closing distance Dand the maximum closing distance D defines the third dimension of theprobability envelope PE.

Thus, referring to FIG. 11, the probability envelope PE at a particularclosing distance D ensures that at least one fragment 355 will impact akill zone KZ that overlaps anywhere with respect to the probabilityenvelope PE. At the same time, the general helical configuration of eachset 370 allows the fragments 355 to be specifically directed towards theprobability envelope and to cover therein a large lateral spread and alarge axial spread of fragments, the fragments originating from thecompact volume provided by the fragmentation warhead 350, that in thisexample is configured for being launched via a Grad type booster stage.

It is also evident that that the fragmentation warhead 300, while beingconfigured for being mounted in the carrier vehicle 102, and having alongitudinal dimension L along a longitudinal axis LA, is alsoconfigured for selectively providing a plurality of fragments 355directed towards a target area spaced at an interception spacing fromfragmentation warhead 300 corresponding to the closing distance D, toprovide a fragmentation pattern on the target area. Referring to FIG.11, for example, the fragmentation pattern includes at least one set,and in this example three sets, of fragments 355 in a spaced linearrelationship, wherein the fragments are spaced at spacing t in each set.The spaced fragments 355 in each set at the target area extend to alength dimension (corresponding to the lateral spread f) greater thanthis longitudinal dimension L, wherein adjacent fragments in each set atthe target area are spaced at a respective fragmentation spacing(spacing t) that is within the dimensions KL or KD of the kill zone KZof the intended target, i.e., the warhead 1. It is further evident thateach fragment 355 is capable of neutralizing the intended target byimpacting the kill zone KZ thereof.

Thus, and referring also to FIG. 6 and FIG. 6A, the fragmentationportion 350 is configured for fragmenting into a plurality of laterallyadjacent sets 370 of serially adjacent fragments 355 in generallyhelical relationship with respect to the longitudinal axis LA, inresponse to detonation of the explosive charge 360. In this example, thefragmentation portion 350 is configured for fragmenting into threelaterally adjacent sets 370 of fragments 355, each set comprisingbetween 30 and 50 serially adjacent fragments 355 in generally helicalrelationship with respect to the longitudinal axis LA, in response todetonation of the explosive charge 360.

It is evident that after detonation of the explosive charge 360, thefragments 355 of each set 370 are serially spaced at a spacing t that isless than a diameter GD of warhead 1 at a distance less than 7.5 m fromthe longitudinal axis LA.

It is also evident that after detonation of the explosive charge 360,each set of fragments 355 is spaced from an adjacent set of fragments355 at a spacing d that is less than the axial length GL of the warhead1 at a distance less than 7.5 m from longitudinal axis LA.

According to an aspect of the presently disclosed subject matter thefragmentation warhead 300 is configured for being maneuvered to apredicted interception point (PIP) via the carrier vehicle 102 forintercepting within a probability envelope a flying object having a killzone of known dimensions.

Turning now to FIG. 12A, this illustrates a generalized detection andtracking scenario that is devoid of seeker head, in accordance withcertain examples of the presently disclosed subject matter. Thus, aflying target rocket 300, i.e., corresponding to the incoming warhead 1of FIG. 8, is detected and tracked by an active sensor 301 (e.g. radaror ladar). Note incidentally that for convenience the term radar ismainly used herein but should be regarded by those skilled in the art asonly an example of active sensor.

The latter emanates radiation which is reflected by the flying targetand received by the active sensor 301 and additional two passivereceiving sensors 302 and 303 whose locations are known in advance atdecimeter level of accuracy. The specified three sensors (constituting asensor array) are configured to utilize respective clocks at a relativesynchronization accuracy of at least 1 nanosecond between each two ofthem and are capable of determining an updated flying rocket's locationat a high accuracy of say less than a meter and therefrom the controlsystem 304 is configured to determine an updated Predicted InterceptionPoint (PIP) in which the interceptor will hit the target. Thepreliminary predicted interception point (PIP), a time of launch andinterceptor missile separation time are transferred as a part of missiondata to the interceptor (not shown) before launching. Updates of PIP aretransferred to the interception during the flight via the uplinkcommunication channel (not shown). The latter requires an estimation ofthe locations of both flying objects (rocket and interceptor) duringtheir flights.

Note that in accordance with certain examples the preliminary missiondata such as predicted PIP missile separation time (loaded to theinterceptor before launch) may be determined based on active sensor onlyand later on during the flight trajectory of the target, say after theascending section of the flight trajectory of the target, the passivesensors can also be utilized allowing determination of more accuratedata such as updated PIP and updated separation time. The utilization ofthe passive sensors may for example depend on the specific threatdefinition and coverage of the passive antennae array.

Note that the determination of updated flying target and interceptorlocations and updated PIP may be performed continuously at a desiredrate, depending on the particular application.

Turning now to FIG. 12B, this illustrates (a possible) ground controlsystem architecture, in accordance with certain examples of thepresently disclosed subject matter; As shown, active sensor 3001 and twopassive sensors 3002 and 3003 are configured to determine the locationof the flying target at high accuracy and send the data to the controlsystem 304. The latter includes the Threat hitting point estimatorsystem 3004 capable of calculating:

-   -   the PIP,    -   the required launch timing of the interceptor missile and    -   the required timing of interceptor missile separation during its        flight.

The Threat hitting point estimator system 3004 transfers the noted abovecalculated parameters to the Interceptor launch manager system 3005capable to generate a mission data that includes at least noted aboveparameters. The Interceptor launch manager transfers the said tolaunching battery 3006, and finally loaded to the interceptor missile,corresponding to the interception missile 100 of FIG. 1. Note that fordetermining the launch timing, the Threat hitting point estimator isalso fed with the sensor's data indicating the precise location of thedetected target rocket. Once the interceptor missile is launched, thethreat hitting point system 3004 is fed also with accurate location dataof the intercepting missile as determined by the ground sensors array.Based on the accurate location data of both the flying target and theinterceptor missile, system 3004 is configured to calculate updated PIPand updated timing of interceptor missile separation and transmit it tothe interceptor missile through uplink communication manager system 3007and the specified data is transmitted through antenna 3008 to theinterceptor missile (and received by uplink receiver module—see, e.g.2007 in FIG. 2C), for further processing by the airborne computeron-board the missile. In addition, the launch time and the planned motorseparation time are also transmitted to the missile.

Reference is now made to FIG. 13 which is a simplified semi-pictorialsemi-functional block diagram illustration of a system for Relative TimeMeasurement between two or more non-collocated stations 410 and 420(e.g. any of the two specified sensors 301 to 303 of FIG. 12A) withdecimeter level accuracy known coordinates, constructed and operative inaccordance with certain examples of the present presently disclosedsubject matter. Each station observes a Common External Signal (e.g.produced by GNSS satellite 430) via antennae 415 and 425 respectively.Each station produces time tag samples (pseudo range and integratedDoppler) based on a common external signal which may be generated by orgenerated responsive to a satellite 430. Each station senses a CommonExternal Event 440 (for example reflection of beam originated by theactive sensor power from the target) and computes a precise Time Periodbetween a sensed external event time tag and the time tag of a latest ofthe samples.

A time offset Computation Unit 450 receives samples from stations A andB and computes a Time Offset between station 410's and 420's clocks atsampling time e.g. using Equations 1-4 below. The time offsetinformation is provided to a nanosecond accuracy processing unit 460which accurately measures time elapsing between events at stations A andB as described in detail below.

The time offset computation performed by unit 460 is (may be based on aconventional Single Difference (SD) algorithm e.g. as described inBradford W. Parkinson and James J. Spilker, Global Positioning System:Theory and applications, Vol. II, Chapter 18, Eq. 9. An instant TimeOffset is computed between the stations 410 and 420's internal timescales using coherent pseudo ranges and integrated Doppler Samples andthe Known Positions of the stations' antennae 415 and 425.

Typically, the Single Difference (SD) algorithm implements the followinglinear combinations of coherent pseudo-range and carrier phase(integrated Doppler), as follows (Equations a and b):

P _(AB) ^(S) =P _(B) ^(S) −P _(A) ^(S)=ρ_(AB) ^(S) +δt _(AB) ·c+B_(AB)+_(I) _(AB) ^(S) +T _(AB) ^(S)+ε^(Code)   (a)

Φ_(AB) ^(S)=Φ_(B) ^(S)−Φ_(A) ^(S)=ρ_(AB) ^(S) +δt _(AB) ·c+B _(AB) −I_(AB) ^(S) +T _(AB) ^(S) +F _(AB) ^(S)+ε^(Phase)   (b)

where samples A provided by station A of FIG. 13 include:

P_(A) ^(S)—Pseudo-range measurement of satellite S (430 in FIG. 13) atstation A, and

Φ_(A) ^(S)—Carrier-phase measurement of satellite S (430 in FIG. 13) atstation A. Samples B provided by station B of FIG. 13 include:

P_(B) ^(S)—Pseudo-range measurement of satellite S (430 in FIG. 13) atstation B, and

Φ_(B) ^(S)—Carrier-phase measurement of satellite S (430 in FIG. 13) atstation B. and wherein:

ρ_(AB) ^(S)—Difference in Ranges between stations A and B and satelliteS (430 in FIG. 13),

C—Speed of light,

B_(AB)—Difference between hardware delays between stations A and B, e.g.as computed by the calibration apparatus of FIG. 13 in PCT applicationno. IL2010/000346 (WO 2010/125569), entitled “Relative Time MeasurementSystem with nanosecond Level Accuracy” (hereinafter “The Time SyncApplication” whose contents are incorporated herein by reference, and asdescribed in detail below

I_(AB) ^(S)—Difference in ionospheric delays between stations A and Band satellite S (430 in FIG. 13),

T_(AB) ^(S)—Difference in tropospheric delays between stations A and Band satellite S (430 in FIG. 13),

F_(AB) ^(S)—Difference in floating ambiguities between stations A and Band satellite S (430 in FIG. 13), e.g. as computed by the calibrationapparatus of FIG. 13 in “The Time Sync Application”,

ε^(Code)—Pseudo-range sampling noise,

ε^(Phase)—Carrier Phase sampling noise,

δt_(AB)—Time difference between stations A and B, e.g. as computed byEquation 5 described below=AB time offset of FIG. 13.

Parameter ρ_(AB) ^(S) is known based on satellite and stations'positions. Parameters I_(AB) ^(S) and T_(AB) ^(S) are modeled usingstandard procedures such as described in the above described textbookGlobal Positioning System: Theory and applications, at Vol. II, Chapter18, Eq. 12, at Vol. I, Chapter 11, Eq. 20, and at Eq. 32. Relative biasB_(AB) is a difference between hardware delays measured once per eachpair of stations. This results in the following equations, which may besolved by the Computation Unit 450 using least squares techniques forunknown Time Offset δt_(AB) and F_(AB) ^(S) respectively (Equations Cand D):

{tilde over (P)} _(AB) ^(S) =δt _(AB) ·c+ε ^(Code)  (c)

{tilde over (Φ)}_(AB) ^(S) =δt _(AB) ·C+F _(AB) ^(S)=ε^(Phase)  (d)

One method of operation for the nanosecond accuracy processing unit 460of FIG. 13 is now described in detail. Based on Time Period which may becomputed by the sensor in stations 410 and 420, e.g. as per Equation Fas described in detail in “The Time Sync Application”, and based also onTime Offset between stations' clocks as derived by Equations C and D,Processing Unit 450 computes a Relative Time Measurement dT_(EVENT)^(AB), also termed herein the “time between events”, between stations410 and 420, e.g. as per following equation E:

dT _(EVENT) ^(AB) =T _(PERIOD) ^(B) −T _(PERIOD) ^(A) +δt _(AB)  (e)

-   -   dT_(EVENT) ^(AB)—Relative Time Measurement of event reception at        stations A and B, also termed “precise relative time” or “time        between events” (FIG. 13),    -   δt_(AB)—Time Offset between station's clocks at sampling time,        typically derived from Equations C and D by Computation Unit 450        as shown in FIG. 13.    -   T_(PERIOD) ^(A)—Time Period between sensing of the external        event (e.g. receiving of a pulse, originated by the active        sensor and reflected by the flying object) by station A as        described in “The Time Sync Application” (equation F). Also        termed (e.g. in FIG. 13) “time period A”,    -   T_(PERIOD) ^(B)—Time Period between sensing of the external        event (e.g. receiving of a pulse, originated by the active        sensor and reflected by the flying object) by station B as        described in “The Time Sync Application” (equation F). Also        termed (e.g. in FIG. 13) “time period B”,

The structure and operation of each station for detecting offset of 1nanosecond is described in detail with reference to FIGS. 2-5 of thespecified “Time Sync Application”.

FIG. 14 shows the less than 1 nanosecond time offset.

Note that the said Relative Time Measurement is performedinstantaneously and continuously at the desired events rate, dependingupon the particular application.

Note also that the station in the specified application refers to asensor of the present application, and that the time offset correspondsto the accuracy of a synchronized clock that is referred to in thepresent application.

Note also that the presently disclosed subject matter is not bound toachieving a time offset of up to 1 nanosecond in accordance with theteachings of “The Time Sync Application”.

Note that the description with reference to FIGS. 13 and 14 describesachieving relative synchronization between two sensors at an accuracy ofup to 1 nanosecond. The specified technique may be applied between anytwo sensors.

Note that the specified technique may be also applied between anypassive sensor and the active sensor (“Star Configuration”. The activesensor clock plays the role of “Master Clock”).

Note also that the specified units 450 and 460 may form part of asynchronization unit forming part of the sensor array and may beimplemented in module or module(s) that are integral with one or more ofthe specified sensors or separated therefrom.

Attention is now drawn to FIG. 15 illustrating a typical interceptionscenario, in accordance with certain examples of the presently disclosedsubject matter and also to FIG. 16 illustrating an exemplary sensorarray layout, in an interception scenario, in accordance with certainexamples of the presently disclosed subject matter. Thus, a targetrocket 601, 701 is detected and tracked at high accuracy utilizing theactive sensor (e.g. radar 702) and the at least two passive sensors (703and 704) (which in certain examples are utilized at a later stage of thetarget's flight, as discussed above), all being clock synchronized at ahigh accuracy of about 1 nanosecond, all as explained in detail above.The accurate updated location determination of the target is based,amongst others, on clock synchronization between the sensors, all aswill be explained in greater detail below. Having accurately determinedthe updated location of the oncoming threat (601, 701) utilizing thespecified sensors, the control system 304 (see FIG. 12A) is capable ofcalculating a planned flight trajectory of the interception missile(designated 602,605,706 in these figures) and updated PredictedInterception Point (PIP) (603,705) for interception the flying threat.The control system 304 (see FIG. 12A) is also capable of establishingthe mission data (not shown), transferred to the interception launcher(not shown) and including at least a preliminary PIP (603), a neededtime of launch and a time of the separation between the interceptionmissile's boost stage 101 and carrier vehicle 102. Thereafter, theinterception missile (602, 605,706) is launched (toward the PIP) and itsupdated location is accurately tracked by the specified sensors (702 to704). As specified, the utilization of the passive sensor give rise to amore accurate mission data (e.g. PIP and separation time).

Note that the control system 304 (see FIG. 12A) can update the PIP(603,705) which is based on the updated locations of the both flyingtarget and the interceptor.

The updated PIP (603) and updated location of the interceptor aretransmitted from the control system 304 (see FIG. 12A) to theinterceptor missile (602, 605,706) at a desired rate depending upon theparticular application.

Note that updated location of the interceptor missile may be used by thenavigation task of interceptor's airborne computer (e.g. 2004 at FIG.3C).

Note also that both updated PIP and the updated location of theinterceptor may be used by guidance task of interceptor's airbornecomputer 2004.

Once the separation condition is achieved: e.g. time of flight ofinterceptor missile meets the separation time condition transmitted fromthe ground control system 304, the booster stage 101 is separated andthe carrier vehicle 102 of the interceptor missile 605 proceeds forward.At this stage the steering fins 107 are extracted and serve steeringsystem for the diverting of the interceptor toward the PredictedInterception Point (PIP). Note that steering towards the PIP should beunderstood as encompassing also steering substantially towards thevicinity of the PIP or hitting it.

The airborne computer 2004 (see FIG. 3C) receives updated navigationdata from the onboard inertial measurement unit and external measurementof interceptor location by uplink communication channel (using thecommunication system 2007—see FIG. 3C). Based on updated navigationsolution and updated PIP (as received by the uplink communication system2007—see FIG. 3C) the airborne computer propagates the guidance law(e.g. proportional navigation) and calculates appropriate steeringcommands for the steering system 2001-2003, see FIG. 3C (by the exampleutilizing steering fins 107 see FIG. 3A) that will guide the interceptormissile toward the updated PIP.

Note that the steering commands are updated based on the updated PIP andupdated location and orientation of the interceptor missile.

Note also that in accordance with certain examples the steering commandsmay be executed after separation of the booster stage 101.

Note that In accordance with certain examples, the interceptorseparation time that forms part of the mission data, may be updated bycontrol system 304 (see FIG. 12A) during the flight of the interceptorvia uplink communication channel 2007 (see FIG. 3C).

In accordance with certain examples, the guidance rules which controlthe steering of the missile towards the PIP may comply (but notnecessarily) e.g. with a known per se proportional navigation paradigm.

Reverting now to FIGS. 15 and 16, once the interception missile (605,701) is sufficiently close to the updated PIP, the proximity fuse (see105 in FIG. 1) of the intercepting missile is activated (in response toa command originated by the airborne computer, or in accordance with acertain other example by the ground control system—see 304 in FIG. 12Aand operates in a known per se manner. Thus, in accordance with certainexamples, the proximity fuse acquires the needed information and once itconcludes that the interceptor missile is sufficiently close to thewarhead of the target, it triggers the fragmentation warhead of theinterceptor missile which detonates the explosive (fitted in thewarhead) and kills the target (601,701).

The target's warhead 1 will be neutralized colliding with and beingpenetrated by the fragments 355 with appropriate mass/densitycharacteristics as disclosed above with reference to FIGS. 1 to 11.

It should be noted that in accordance with certain examples due to therelatively small dimensions of the voluntary area of the GTG rocket, theminimal required fragments' density is about one lethal fragment persquare decimeter. Reliable interception of the target with a very smallvulnerable area by an interceptor can be achieved only by very accurateestimation of the interception point (required miss distance issub-meter level).

Note that the presently disclosed subject matter is not bound by thespecified operational specifications which are provided for illustrativepurposes only, and the latter may vary depending upon the particularapplication.

Note that using other MLRS solid motors for acceleration of theinterceptor provides a capability for carrying of sufficiently largerwarhead.

Note also that in accordance with certain examples any known per selethality criteria for activating the warhead may be employed. Note thatthe lethality criterion may vary depending upon the nature of thetarget. Say, for rocket a different lethality criterion may be employedcompared for instance to a UAV or aircraft.

Turning now to FIG. 17, it illustrates a deployment scheme of combined(radar and multi-static) sensor synchronized at a high accuracy of about1 nanosecond, in accordance with certain examples of the presentlydisclosed subject matter.

Before turning to FIG. 17, it should be noted that Target location ismeasured by three independent highly accurate measurements:

-   -   Range (R_(t)) from Radar's location to target and is measured by        the radar itself (at decimeter level),    -   Time difference (Δt₁) the time difference between the reception        times by the radar and antenna #1 of a signal that is emanated        by the radar and reflected from the target. Note that the radar        is located at origin of the coordinate system 801), the target        is located at (X_(t), Y_(t), Z_(t)) 804 and antenna #1 is        located at (0, Y_(ant1), 0) 802. The time difference is measured        by synchronization system (1 nanosecond level, equivalent to        decimeter level measurement).    -   Time difference (Δt₂) the time difference between the reception        times by the radar and antenna #2 of a signal that emanated by        the radar and reflected from the target. Note that antenna #2 is        located at (X_(ant2), Y_(ant2), 0) 803.

The instantaneous target location is determined by using the known perse Time Difference Of signal Arrival (TDOA) technique (see for exampleTDOA Localization Techniques IEEE 802.15-04a/572r0 ieee802.org/ . . ./15-04-0572-00-004a-tdoa-localization-techniques.ppt, October 2004),combined with the measurements of range to target. The target locationmeasurement uncertainty volume is an intersection of:

-   -   Spherical layer with radius of R_(t) and thickness equal to        uncertainty of range measurement    -   Hyperbolical layer Δt₁=constant and thickness equal to product        of light velocity and uncertainty of synchronization of        reflection power receiving by the radar and the passive antenna        #1.    -   Hyperbolical layer Δt₂=constant and thickness equal to product        of light velocity and uncertainty of synchronization of        reflection power receiving by the radar and the passive antenna        #2.

Accuracy of target trajectory estimation may be approved byimplementation of e.g. known per se filtering procedure (Kalman Filter).This method uses the multiple results of target location measurementsand predefined models of target kinematic behavior.

Accuracy of instantaneous measurement may be approved by using of e.g.known per se TDOA-FDOA technique (noted above TDOA technique combinedwith Frequency Difference Of signal Arrival technique) that requiredadditional measurements of frequency shifting by the Doppler Effect.Effectiveness of noted improvement depends on:

-   -   Stability of frequency sources used at radar and passive        antennae sites,    -   Radar signal type and signal processing (resolution of Range        Doppler map).    -   Target kinematic characteristics (velocity, spin rate, Radar        Cross Section pattern etc.).

The form of the noted above uncertainty volume of target locationdepends on deployment of the radar and the passive antennae. The knownper se technique of Geometric Dilution Of Precision—GDOP may be used (asan example) for optimization of divided sensor deployment (see forexample Richard B. Langley (May 1999). “Dilution of Precision”.GPSWorld. http://gauss.gge.unb.ca/papers.pdf/gpsworld.may99.pdf.

The special case of the plane deployment of the divided sensor isanalyzed below. In this case all parts of the sensor array (radar andpair of passive antennae) are ground based and as a result theelongation of Vertical Dilution Of Precision (VDOP) is expected.

A method for calculation of a Y-coordinate of the target is illustratedby FIGS. 18, 18A and 18B. The noted above spherical layer with radius R(904) is generated by a range measurement by radar A (901) placed at ahead of a coordinate system XYZ. The hyperbolic layer Δt₁=constant (905)is a place of points with the constant difference of the distances tothe radar A (901) and a passive antenna C (902). The intersectionbetween the noted sphere and hyperbolic layers generates a ring (906) ofpossible positions of a target B (903). Turning now to FIG. 18A, atriangle ABC is generated by a radar site A (901), 1st passive antennasite C (902) and a target B (903). The length of side AB is equal torange R (904), accurately measured by the radar, the length of side ACis equal to “Y” coordinate of 1-st passive antenna site Y_(ant1) and thelength of side BC (905) is equal to a distance L₁ of a traveling path ofreflected by the target RF energy, originated by the radar. An angle φ(906) is an angle BAC between side AB and AC (Y-axes). A point B′ (907)is a projection of the target B on the XY-plane. Equation (1) presents atrigonometric relationship (cosines law) for ABC triangle:

R ² +Y _(ant1) ²−2·R·Y _(ant1) cos(φ)=L ₁ ²  (1)

The term cos (φ) can be calculated by ratio

$\frac{{AB}^{''}}{R},$

here B″ (908) is projection of point B on the Y axes (AB″=Y_(T)).

R ² +Y _(ant1) ²−2·Y _(ant1) Y _(T) =L ₁ ²  (2)

The distances R and R-L₁ are measured accurately by the sensor array.The noted above time difference (Δt₁) between receiving of the reflectedby the target RF energy by the radar and the passive antenna #1 isconnected to the distance R-L₁ by following relationship:

R−L ₁ =c·Δt ₁, or L ₁ =R−c·Δt ₁,  (3)

-   -   Here c is a speed of light

Substituting of obtained terms for L₁ (equation 3) into relationship (2)provides the following equation for a Y-coordinate of a target:

$\begin{matrix}{Y_{T} = {\frac{Y_{{ant}\mspace{11mu} 1}^{2} - {c^{2}\Delta \; t_{1}^{2}}}{2 \cdot Y_{{ant}\mspace{11mu} 1}} + {{\frac{R}{Y_{{ant}\mspace{11mu} 1}} \cdot c \cdot \Delta}\; t_{1}}}} & (4)\end{matrix}$

Note that for obtaining a Y coordinate of the target Y_(T) only 2accurate measurements of the sensor array were used: range to target Rand time difference between receiving of reflected by target RF energyby the radar and by the first passive antenna (Δt₁).

First order approximation of an accuracy of target's Y coordinate (ε_(Y)_(t) ) determination is presented by the following equation:

$\begin{matrix}{{\left. ɛ_{Y_{t}} \right.\sim\frac{c}{Y_{{ant}\mspace{11mu} 1}}} \cdot \sqrt{\left( ɛ_{t \cdot L_{1}} \right)^{2} + \left( {{ɛ_{R} \cdot \Delta}\; t_{1}} \right)^{2}}} & (5)\end{matrix}$

Here ε_(t) and ε_(R) are independent accuracies of the measurements oftime difference and a range correspondingly.

The following rule provides a capability for a decimeter level accuracydetermination of the Y coordinate of the target: the first passiveantenna should be placed close to the area of potential interceptionpoints (L₁˜Y_(ant1)) and relatively far from the radar site(Y_(ant1)˜R_(t)).

FIG. 18B illustrates a dependency of the accuracy of determination ofthe Y coordinate of the target on the main parameters of the radar andpassive antenna deployment (L₁/Y_(ant1) and R_(t)/Y_(ant1)). The notedabove accuracy is kept on the level of 0.5-0.6 in the case ofL₁/Y_(ant1)<1.5 & R_(t)/Y_(ant1)<1.5.

Note that the discussion with reference to FIGS. 17 and 18, 18A, 18B isprovided for illustrative purposes and accordingly those versed in theart will readily appreciated that various numerical and specificparameters that were described are by no means binding.

FIG. 19 illustrates a method for calculation X-coordinate of the target.A triangle ABD is generated by a radar site 1001, 2-nd passive antennasite 1002 and a target 1003. The length of site AB is equal to range R(1004), accurately measured by the radar, the length of site AD is equalto distance of 2-nd passive antenna site from the radar site √{squareroot over ((X_(ant2) ²+Y_(ant2) ²)} and the length of site DC (1005) isequal to a distance L₂ of a traveling path of reflected by the target RFenergy emanated by the active sensor 1001. A point B′ (1007) is aprojection of the target B on the XY-plane. Equation (6) presents a wellknown trigonometric relationship (cosines law) for ABD triangle:

R ² +X _(ant2) ² +Y _(ant2) ²−2·R√{square root over ((X _(ant2) ² +Y_(ant2) ²)}·cos(ψ)=L ₂ ²  (6)

The term cos (ψ) can be calculated by ratio

$\frac{AE}{R},$

here E (1008) is projection of point B′ to the site AD:

$\begin{matrix}{{AE} = \frac{{X_{t} \cdot X_{{ant}\mspace{11mu} 2}} + {Y_{t} \cdot Y_{{ant}\mspace{11mu} 2}}}{\sqrt{X_{{ant}\mspace{11mu} 2}^{2} + Y_{{ant}\mspace{11mu} 2}^{2}}}} & (7)\end{matrix}$

Equation (8) presents the relationship for calculation the X coordinateof the target:

$\begin{matrix}{X_{T} = {\frac{X_{{ant}\mspace{11mu} 2}^{2} + Y_{{ant}\mspace{11mu} 2}^{2} - {{c^{2} \cdot \Delta}\; t_{2}^{2}}}{2 \cdot X_{{ant}\mspace{11mu} 2}} + \frac{{R \cdot c \cdot \Delta}\; t_{2}}{X_{{ant}\mspace{11mu} 2}} - \frac{Y_{t} \cdot Y_{{ant}\mspace{11mu} 2}}{X_{{ant}\mspace{11mu} 2}}}} & (8)\end{matrix}$

The range R and the time difference Δt₂ are measured accurately by thesensor array. The Y coordinate of the target Y_(t) is determined byequation (4).

Note that for obtaining of a X coordinate of the target X_(T) all 3accurate measurements of the divided ground sensor were used:

-   -   Range to target R,    -   Time difference between receiving of reflected by the target RF        energy by the radar and by the first passive antenna (Δt₁) for        obtaining the Y coordinate of the target.    -   Time difference between receiving of reflected by the target RF        energy by the radar and by the second passive antenna (Δt₂).

First order approximation of an accuracy of target's X coordinate (Δ_(X)_(t) ) determination is presented by following equation:

$\begin{matrix}{\left. ɛ_{X_{t}} \right.\sim\sqrt{{{\frac{c^{2}}{X_{{ant}\mspace{11mu} 2}^{2}} \cdot \left( {{ɛ_{t}^{2} \cdot L_{2}^{2}} + {{ɛ_{R}^{2} \cdot \Delta}\; t_{2}^{2}}} \right)} + {\frac{Y_{{ant}\mspace{11mu} 2}^{2}}{X_{{ant}\mspace{11mu} 2}^{2}} \cdot ɛ_{Y_{t}}^{2}}},}} & (9)\end{matrix}$

Here ε_(t) and ε_(R) are accuracies of the measurements of timedifference and range correspondingly.

The following rule provides capability for a decimeter level accuracydetermination of the X coordinate of the target: the second passiveantenna should be placed close to X-axes of the chosen coordinate system(Y_(ant2)<<X_(ant2)). In this case the relationship for theuncertainties of determining of X coordinate of the target has thefollowing form:

$\begin{matrix}{{{\left. ɛ_{X_{t}} \right.\sim\frac{c}{X_{{ant}\mspace{11mu} 2}}} \cdot \sqrt{{ɛ_{t}^{2} \cdot L_{2}^{2}} + {{ɛ_{R}^{2} \cdot \Delta}\; t_{2}^{2}}}},} & (10)\end{matrix}$

In the case of orthogonal deployment of the combined radar and multistatic array (Y_(ant2)=0), the expression for the accuracy ofdetermination of the X coordinate of the target (10) is similar to theexpression for the accuracy of determination of the Y coordinate of thetarget (5). In accordance with certain examples, the following ruleprovides a capability for a decimeter level accuracy determination ofthe X coordinate of the target:

-   -   the second passive antenna should be placed relatively far from        the radar site (X_(ant2)˜R_(t)) and relatively close to the area        of potential interception points    -   (X_(ant2)˜L₂). Dependency of the accuracy of determination of X        coordinate of the target on the main parameters of second        passive antenna deployment is similar to the dependency of the        accuracy of Y coordinate: the noted above accuracy is kept on        the level of 0.5-0.6 in the case of L₂/Y_(ant2)<1.5 &        R_(t)/Y_(ant2)<1.5 (see FIG. 9B).    -   R_(t)/X_(ant2), L₂/X_(ant2) In the case of non-orthogonal        deployment of the combined radar and multi static array, the        expression for the accuracy of determination of the X coordinate        of the target (9) includes additional term (Y_(ant2) ²/X_(ant2)        ²·ε_(Y) _(t) ²). This term decreases the accuracy of        determination of the X coordinate of the target. For example, in        the case of deployment of the passive antennae with the angle of        60° in respect to the radar) (Y_(ant2) ²/X_(ant2) ²=(1/tan        60°)²=⅓), the accuracy of determination of X coordinate of the        target will be about 15%    -   √{square root over (1.333)}−1, worse relatively to the optimal        orthogonal deployment.

Substituting of obtained terms for Y_(t) (equation 5) and X_(t)(equation 8) coordinates of the target into the equation for the rangeallows calculation of the last target coordinate (Z_(t)):

$\begin{matrix}{Z_{t} = {\sqrt{R_{t}^{2} - Y_{t}^{2} - X_{t}^{2}} = \sqrt{\begin{matrix}{R_{t}^{2} - \left( \frac{{\frac{1}{2} \cdot \left( {X_{{ant}\mspace{11mu} 2}^{2} + Y_{{ant}\mspace{11mu} 2}^{2} - {{c^{2} \cdot \Delta}\; t_{2}^{2}}} \right)} + {{R \cdot c \cdot \Delta}\; t_{2}} - {Y_{t} \cdot Y_{{ant}\mspace{11mu} 2}}}{X_{{ant}\mspace{11mu} 2}} \right)^{2} -} \\\left( \frac{{\frac{1}{2} \cdot \left( {Y_{{ant}\mspace{11mu} 1}^{2} - {{c^{2} \cdot \Delta}\; t_{1}^{2}}} \right)} + {{R \cdot c \cdot \Delta}\; t_{1}}}{Y_{{ant}\mspace{11mu} 1}} \right)^{2}\end{matrix}}}} & (11)\end{matrix}$

If the sensor array is deployed according to the formulated above rules:Y_(ant1)˜R_(t), X_(ant2)˜R_(t), Y_(ant2)<<Y_(ant1), the equation forZ-coordinate of the target has the following form:

$\begin{matrix}{Z_{t} = \sqrt{\begin{matrix}{R_{t}^{2} - \left( \frac{{\frac{1}{2} \cdot \left( {X_{{ant}\mspace{11mu} 2}^{2} - {{c^{2} \cdot \Delta}\; t_{2}^{2}}} \right)} + {{R \cdot c \cdot \Delta}\; t_{2}}}{X_{{ant}\mspace{11mu} 2}} \right)^{2} -} \\\left( \frac{{\frac{1}{2} \cdot \left( {Y_{{ant}\mspace{11mu} 1}^{2} - {{c^{2} \cdot \Delta}\; t_{1}^{2}}} \right)} + {{R \cdot c \cdot \Delta}\; t_{1}}}{Y_{{ant}\mspace{11mu} 1}} \right)^{2}\end{matrix},}} & \left( 11^{\prime} \right)\end{matrix}$

An accuracy of determination Z coordinate of the target can be estimatedby the following equation:

$\begin{matrix}\begin{matrix}{ɛ_{Z} = \frac{Z_{\max}^{2} - Z_{\min}^{2}}{2 \cdot Z_{t}}} \\{= \frac{\begin{matrix}{\left( {R_{\max}^{2} - X_{\min}^{2} - Y_{\min}^{2}} \right) -} \\\left( {R_{\min}^{2} - X_{\max}^{2} - Y_{\max}^{2}} \right)\end{matrix}}{2 \cdot Z_{t}}} \\{{= {{\frac{R_{t}}{Z_{t}} \cdot ɛ_{R}} + {\frac{X_{t}}{Z_{t}} \cdot ɛ_{X}} + {\frac{Y}{Z_{t}} \cdot ɛ_{Y}}}},}\end{matrix} & (12)\end{matrix}$

Accuracy of the target's Z-coordinate determination improves with theincreasing of the targets altitude (Z_(t)), that is why an interceptionof the target close to its apogee is preferable. High accelerationduring the interceptor missile boost phase can significantly improve asystem time budget. Rocket solid motors (for example, the motor of GRADor MLRS rockets) are usually designed for extremely short burning timeand can be useful as a low cost propulsion part of the interceptor.

Equation 13 outlines a different form of equation 12:

$\begin{matrix}{ɛ_{Z} = {{\frac{R_{t}}{Z_{t}} \cdot ɛ_{R}} + {\frac{X_{t}}{Z_{t}} \cdot ɛ_{X}} + {\frac{Y_{t}}{Z_{t}} \cdot ɛ_{Y}}}} & (13)\end{matrix}$

What remains to finalize the deployment of the sensor's array layout isthe location of the active sensor (e.g. radar).

As before, it is desired to reduce ε_(z) (see equation 13) in order tosecure hitting the target. Before moving on, it is recalled that shortlyafter the detection of the flying GRAD threat, the interceptor missile(e.g. powered by GRAD or MLRS motor) is launched towards the target froma launching site. Both fly at substantially the same speed andsubstantially along known trajectories, which substantially prescribethe predicted interception point.

Reverting to equation 13, the lower the expression R_(t)/Z_(t), thelarger the ε_(z) (the other variables including Z_(t) are substantiallyknown). This stipulates that the target range R_(t) should be smaller.Assuming that by certain examples the radar cannot view backwardly, thenthe most advantageous location would be substantially underneath thepredicted interception point. In certain examples, the radar can viewbackwardly, implying thus that it can be deployed farther than the PIP.

In an exemplary interception scenario the target acquisition starts atthe relatively low ascent part of target trajectory (elevation angle ofthe target is less 20 degrees, R_(t)/Z_(t)˜3÷10). The contribution ofthe passive antennae array to the improving of the target locationaccuracy is limited by factor R_(t)/Z_(t). It is sufficient for thedefinition of PIP, generation of mission data and launching of theinterceptor towards the PIP but not enough for meeting the lethalitycriterion for destroying the target (e.g. hitting the target). Along thetarget trajectory the elevation angle of radar beam rises up and e.g. atthe apogee of the target can reach about 45 degrees (R_(t)/Z_(t)˜1.4).As a result, the accuracy of the measurement of the target locationsignificantly improves: the expected accuracies of determination of Xand Y coordinates of the target location (according to the equations 5and 10) are close to the 0.5÷1 meter and of Z coordinate of the target(according to the equation 13) is close to 2÷2.5 meter. At the descentpart of the target trajectory the elevation angle of the radar beamincreases continuously and can reach e.g. about 60 degrees at the regionof potential interception points (R_(t)/Z_(t)˜1.15). The measurements ofX and Y coordinates remain to be very accurate (0.5÷1 meter level) andaccuracy of measurement of Z coordinate of the target reaches level of1.5÷2 meter. The volume of uncertainty of target location is smallenough (1÷2 m³) for secure hitting of the target warhead by the beam offragments generated by the interceptor with relatively small warhead.Note that the presently disclosed subject matter is not bound by thespecified exemplary scenario and in particular not by the specificnumerical parameters outlined in the scenario.

The net effect is this that optimal deployment in accordance withcertain examples of the presently disclosed subject matter stipulatesthat the first passive antenna will be deployed in the direction thatfalls in the sector from which the oncoming threat is likely to arriveand at coordinates (0,Y_(ant1),0). The direction is from the activesensor to the first antenna. The second antenna will be deployed closeto perpendicular direction e.g. at coordinates (X_(ant2), Y_(ant2),0,)where Y_(ant2)<<Y_(ant1) and that the radar will be placed as far aspossible, preferably (in the case of a radar that is devoid of backwardview) underneath the predicted interception point (and in the case ofbackwardly viewing radar, further than the PIP) in order to decrease therange to target R and that Y_(ant1)˜R_(t) and X_(ant2)˜R_(t). Note that“˜” is indicative of up to say 1.5 times, e.g. if the range to targetfrom the radar site is 15 Km than the distances Y_(ant1) and X_(ant2)could be at least 10 Km.

Note that the discussion with reference to FIG. 19 is provided forillustrative purposes and accordingly those versed in the art willreadily appreciated that various numerical and specific parameters thatwere described are by no means binding.

Turning now to FIG. 20 it illustrates a sequence of operations, inaccordance with certain examples of the presently disclosed subjectmatter.

Thus, at stage 1101, the radar sensor detects and tracks the flyingrocket.

At stage 1102 the control system determines a preliminary inaccuratePredicted Interception Point.

At stage 1103 said control system commands to launch the interceptormissile and transfers to the interceptor the mission data that includesat least required time of launch, inaccurate PIP and required time ofinterceptor separation. The radar sensor continues to track the flyingrocket.

At stage 1104 the radar simultaneously continues to track the flyingrocket, detects and starts to track the interceptor missile. The radarsensor transfers the measurement data (tracks) to the said controlsystem. The control system calculates the updated the target rocket andinterceptor missile state vectors (locations and velocities) as well asupdated predicted interception point.

At stage 1105 the interceptor performs the separation and the engine isdiscarded. The interceptor's fins start control the roll of theinterceptor's main section and stop the interceptor spinning.Interceptor's uplink receiver is ready for communication.

At stage 1106 the radar continues to track both flying objects (targetand interceptor) and transfers the measurement data to the controlsystem. The control system continues updating the state vectors of thetarget and interceptor and calculates the updated PIP.

At stage 1107 the radar and the control system continue the operationsof the previous stage, and, in addition, the control unit transmits theupdated data to the interceptor via uplink communication channel. Theuplink message includes at least updated location of the interceptormissile (using by navigation task of the airborne computer of theinterceptor missile) and updated PIP (using by the guidance task of theairborne computer of the interception missile).

At stage 1108 the radar and the control system continue operation of theprevious stage. The interceptor receives the uplink message and uses itby implementing navigation and guidance tasks.

At stage 1109 the radar, the control system and the interceptor continueoperation of the previous stage and in addition the control task ofinterceptor provides steering commands to the steering system. Thesteering system guides the interceptor toward the updated PIP.

At stage 1110 the radar, the control system and the said interceptorcontinue operating of the previous stage and in addition the two otherpassive sensors (synchronized clocks at an accuracy of at least 1nanosecond) receive the reflections of beam originated by the radar. Thetime differences between receiving the echo by the radar, and by thepassive sensors, is transferred to the control system.

At stage 1111 the radar, the control system, the interceptor and thepassive sensors continue operating the previous stage and in additionthe control unit calculates the updated PIP and updated interceptionlocation at a high accuracy.

At stage 1112 the radar, the control system, the interceptor and thepassive sensors continue operating the previous stage and in additionthe control unit transmits the updated PIP and updated interceptionlocation to the interceptor via the uplink communication channel.

At stage 1113 the radar, the control system, the interceptor and thepassive sensors continue operating the previous stage and in additionthe interceptor calculates the timing for the proximity fuse activation.

At stage 1114 in vicinity of said PIP the proximity fuse achieves thecondition for the warhead activation (when a hitting condition is met).

At stage 1115 in vicinity of said PIP, the explosive of theinterceptor's warhead is detonated, and afterwards the lethal fragmentsof interceptor's warhead hit the warhead of the target and neutralizeit.

Note that the system architecture FIGS. 3C, 12B, 13, are provided forillustrative purposes only and are by no means biding. Accordingly, thesystem architecture of each of the specified drawings may be modified byconsolidating two or more blocks/modules/units/systems and/or bymodifying at least one of them and or by deleting at least one of themand replacing by one or more others, all as required and appropriate,depending upon the particular implementation.

Note that the flow chart illustrating sequence of operation in FIG. 20is provided for illustrative purposes only and is by no means biding.Accordingly, the operational stages may be modified by consolidating twoor more stages and/or by modifying at least one of them and or bydeleting at least one of them and replacing by one or more others, allas required and appropriate, depending upon the particularimplementation.

Unless specifically stated otherwise, as apparent from the followingdiscussions, it is appreciated that throughout the specificationdiscussions utilizing terms such as “achieving”, “generating”,“updating”, utilizing” and “activating” or the like, include actionsand/or processes of a computer that manipulate and/or transform datainto other data, said data represented as physical quantities, e.g. suchas electronic quantities, and/or said data representing the physicalobjects. The term “computer” should be expansively construed to coverany kind of electronic device with data processing capabilities,including, by way of non-limiting example, a personal computer, aserver, a computing system, a communication device, a processor (e.g.digital signal processor (DSP), a microcontroller, a field programmablegate array (FPGA), an application specific integrated circuit (ASIC),etc.), any other electronic computing device, and/or any combinationthereof.

The operations in accordance with the teachings herein may be performedby a computer specially constructed for the desired purposes or by ageneral purpose computer specially configured for the desired purpose bya computer program stored in a computer readable storage medium.

Alternatively, any suitable target detection and tracking system can beused for providing the carrier vehicle 102 with data relating to the PIPof the flying object, and relating to the state vectors of the carriervehicle and of the flying object, to enable guiding of the carriervehicle 102 to the PIP. For example, such target detection and trackingsystem can include, in other alternative variations of the aboveexamples, an on-board homing sensor (e.g. RF seeker, electro opticalsensors, and so on) that autonomously detect and track the flyingobject.

It is also to be noted that the fragmentation warhead and carriervehicle according to examples of the presently disclosed subject mattercan also be used for different targets, for example different types orrockets or missiles, UAV's, manned aircraft, cruise missiles, and so on,as well as non-flying objects, for example ground vehicles or marinevehicles, or static structures, such as for example Radar antennas andso on. For example, the closely spaced fragments provided by each set offragments after ejection from the fragmentation warhead can providesevere weakening of a mechanical structure along a particular direction,which can result in the failing or collapse of the structure due tomechanical or aerodynamic loads on the weakened structure.

It is also to be noted that in another alternative variations of theabove examples, the interception missile can be configured for beingair-launched, for example by a carrier aircraft.

In the method claims that follow, alphanumeric characters and Romannumerals used to designate claim steps are provided for convenience onlyand do not necessarily imply any particular order of performing thesteps.

It should be noted that the word “comprising” as used throughout theappended claims is to be interpreted to mean “included but not limitedto”.

Whilst some particular examples have been described and illustrated withreference to some particular drawings, the artisan will appreciate thatmany variations are possible which do not depart from the general scopeof the presently disclosed subject matter, mutatis mutandis.

1. A fragmentation warhead configured for being mounted in a carriervehicle, the warhead having a longitudinal axis and comprising: a shellextending along said longitudinal axis and comprising a fixed shellportion and a fragmentation portion, and defining therebetween a cavityfor accommodating therein an explosive charge; the fragmentation portioncomprising at least one set of serially adjacent fragments incorrespondingly serially contiguous relationship in the fragmentationportion and in generally helical relationship with respect to thelongitudinal axis.
 2. The fragmentation warhead according to claim 1,wherein the fragmentation portion is configured for fragmenting intosaid at least one set of serially adjacent fragments in generallyhelical relationship with respect to the longitudinal axis, in responseto detonation of the explosive charge.
 3. A fragmentation warheadconfigured for being mounted in a carrier vehicle, the warhead having alongitudinal axis and comprising: a shell extending along saidlongitudinal axis and comprising a fixed shell portion and afragmentation portion, and defining therebetween a cavity foraccommodating therein an explosive charge; the fragmentation portionconfigured for fragmenting into at least one set of serially adjacentfragments in generally helical relationship with respect to thelongitudinal axis, in response to detonation of the explosive charge. 4.The fragmentation warhead according to claim 3, wherein prior to saiddetonation, said at least one set of serially adjacent fragments arecorrespondingly serially contiguous relationship in said fragmentationportion and in generally helical relationship with respect to thelongitudinal axis.
 5. The fragmentation warhead according to any one ofclaims 1 to 4, wherein the fragmentation portion is configured forfragmenting into a plurality of laterally adjacent said sets of seriallyadjacent fragments in generally helical relationship with respect to thelongitudinal axis, in response to detonation of the explosive charge. 6.The fragmentation warhead according to any one of claims 1 to 5, whereinthe fragmentation portion is configured for fragmenting into threelaterally adjacent said sets, each said set comprising between 30 and 50said serially adjacent fragments in generally helical relationship withrespect to the longitudinal axis, in response to detonation of theexplosive charge.
 7. The fragmentation warhead according to any one ofclaims 1 to 6 wherein said fixed shell portion is configured so thatupon initiation of detonation of the explosive charge, shockwavespropagating therefrom are directed via said fixed shell portion thetowards said fragmentation portion.
 8. The fragmentation warheadaccording to any one of claims 1 to 7, wherein said fixed shell portionhas rotational symmetry about said longitudinal axis.
 9. Thefragmentation warhead according to any one of claims 1 to 8, whereinsaid fixed shell portion has a generally tubular configuration.
 10. Thefragmentation warhead according to any one of claims 1 to 9, whereineach said set of serially adjacent fragments in correspondingly seriallycontiguous relationship in said fragmentation portion and in generallyhelical relationship with respect to the longitudinal axis is orientedat predetermined helix angle with respect to said longitudinal axis. 11.The fragmentation warhead according to claim 10, wherein said helixangle is predetermined such that upon said detonation, the respectivesaid fragments of each said set are spread over an imaginary cylindricalsurface along a distance of between about 2 m to about 4 m over saidcylindrical surface, at a corresponding radial distance of between 4 mand 8 m, respectively, from said longitudinal axis, while ensuring aspacing of not greater than 0.1 m between adjacent fragments at saidradial distance.
 12. The fragmentation warhead according to any one ofclaims 10 to 11, wherein said helix angle is between 2.5° and 3°. 13.The fragmentation warhead according to any one of claims 1 to 12,wherein said fragmentation portion is formed as a plurality of axiallyserially adjacent fragmentation portion sections, each saidfragmentation portion section comprising a plurality of said fragmentsin lateral contiguous (abutting) relationship, and wherein saidplurality of fragments of each successive said fragmentation portionsection along said longitudinal axis is angularly displaced about thelongitudinal axis with respect to the respective said plurality offragments of the previous said fragmentation portion section.
 14. Thefragmentation warhead according to claim 13, wherein a respective saidplurality of fragments of a first said fragmentation portion section atone longitudinal end of said fragmentation portion is angularlydisplaced about the longitudinal axis with respect to a respective saidplurality of fragments of a last said fragmentation portion section atanother longitudinal end of said fragmentation portion by an angulardisplacement of between 25° and 35°.
 15. The fragmentation warheadaccording to claim 13 or claim 14, wherein the respective said pluralityof fragments of each pair of successive said fragmentation portionsections are angularly displaced from one another with reference to(along) a plane orthogonal to the longitudinal axis by a fragmentationset dispersion angle.
 16. The fragmentation warhead according to claim15, wherein said fragmentation set dispersion angle is between 0.5° and0.7°.
 17. The fragmentation warhead according to any one of claims 1 to16, wherein said fragmentation portion is formed as a generally helicalband with respect to said longitudinal axis, and wherein said fixedshell portion comprises a generally helical slot complementary to saidhelical band.
 18. The fragmentation warhead according to claim 17,wherein said generally helical band is projectable to a plane to providea two dimensional parallelogram pattern of said fragments.
 19. Thefragmentation warhead according to claim 18, said parallelogram patternhaving a base corresponding to the width of three said fragments, and aheight corresponding to the axial length of the fragmentation warhead.20. The fragmentation warhead according to claim 18 or claim 19, whereineach said fragment has a weight of between 25 g and 35 g.
 21. Thefragmentation warhead according to any one of claims 1 to 20, whereinsaid fragmentation portion contains a total number of said fragmentsbetween 100 said fragments and 200 said fragments.
 22. The fragmentationwarhead according to any one of claims 1 to 21, wherein each saidfragment has a plan shape in the form of a parallelogram.
 23. Thefragmentation warhead according to any one of claims 1 to 22, whereinfollowing detonation of the explosive charge, said fragments of eachsaid set are serially spaced at a spacing less than 0.1 m at a distancebetween 4.5 m and 7.5 m from said longitudinal axis.
 24. Thefragmentation warhead according to any one of claims 1 to 23, whereinfollowing detonation of the explosive charge, each said set of fragmentsis spaced from an adjacent said set of fragments at a spacing less than0.50 m at a distance between 4.5 m and 7.5 m from said longitudinalaxis.
 25. The fragmentation warhead according to any one of claims 1 to24, wherein each said fragment is capable of neutralizing a flyingobject by impacting a kill zone thereof.
 26. The fragmentation warheadaccording to claim 25, wherein said kill zone has a length of 0.50 m anda width of 0.10 m.
 27. The fragmentation warhead according to claim 25or claim 26, wherein said flying object is any one of: a rocket, a GRADrocket, a UAV, a manned aircraft, a cruise missile.
 28. A carriervehicle for a fragmentation warhead, comprising: the fragmentationwarhead as defined in any one of claims 1 to 27; an uplink for receivingcommands from a control center; a proximity fuse operatively connectedto the fragmentation warhead and configured for detonating the warheadat a predetermined spacing between the carrier vehicle and a flyingobject; the carrier vehicle being maneuverable at least responsive toreceiving said commands.
 29. The carrier vehicle according to claim 28,wherein the carrier vehicle is configured for being mounted in a boosterstage.
 30. The carrier vehicle according to any one of claims 28 to 29,wherein said proximity fuse is configured for generating two flat laserbeams and for fusion time determination based on reflections receivedfrom said beams.
 31. The carrier vehicle according to any one of claims28 to 30, wherein said uplink comprises a receiver for receiving data orsignals relating to PIP, target and carrier vehicle state vectors,and/or relative state vectors between target and carrier vehicle. 32.The carrier vehicle according to any one of claims 28 to 31, comprisinga plurality of pivotable vanes for steering (maneuvering) said carriervehicle.
 33. The carrier vehicle according to any one of claims 28 to32, further comprising a homing sensor, configured for autonomouslyhoming onto a target.
 34. A missile for intercepting a flying object,comprising: (a) a carrier vehicle for a fragmentation warhead as definedin any one of claims 28 to 33; and (b) a booster stage for propellingthe carrier vehicle along a desired trajectory.
 35. The missileaccording to claim 34, wherein said booster stage is based on a GRADrocket system or wherein said booster stage comprises a GRAD rocketmotor.
 36. An interception system comprising: a missile batteryincluding at least one missile for intercepting a flying object asdefined in any one of claims 34 to 35; a radar system for detecting andtracking at least one said flying object; a control center fordetermining a predicted impact point (PIP) for the missile; acommunications uplink to provide maneuvering data to the carrier vehicleduring flight thereof to intercept the respective said flying object atthe respective predicted interception point PIP.
 37. The systemaccording to claim 36, configured for causing the carrier vehicle to beselectively oriented at a desired relative angle to a flight path of theflying object at the predicted interception point (PIP).
 38. The systemaccording to claim 37, wherein said relative angle is between 10° and12° for a pursuit interception scenario or between 190° and 192° for ahead-on interception scenario, or wherein said relative angle is between−20° to +40° for said pursuit interception scenario, or between +160° to+220° for said head-on interception scenario.
 39. A method forintercepting a flying object, comprising: (i) providing a missile asdefined in any one of claims 34 to 35; (ii) using the booster stage toselectively launch the carrier vehicle along an intercept trajectorywith respect to the flying object; (iii) maneuvering the carrier vehicleinto proximity with the flying object; (iv) detecting the flying objectwithin a minimum range with respect to the fragmentation warhead via theproximity fuse; and (vi) detonating the explosive charge responsive tostep (iii).
 40. The method according to claim 39, wherein in step (iii)the carrier vehicle is oriented at a relative angle to the flying objectat a predicted interception point (PIP).
 41. The method according toclaim 40, wherein said relative angle is between 10° and 12° for apursuit interception scenario or between 190° and 192° for a head-oninterception scenario, or wherein said relative angle is between −20° to+40° for said pursuit interception scenario, or between +160° to +220°for said head-on interception scenario.
 42. The method according to anyone of claims 40 to 41, wherein to provide a spacing between the carriervehicle and the flying object of between 4.5 m and 7.5 m at the PIP,and/or wherein the fragmentation portion is facing the flying object atthe PIP.
 43. An interception missile comprising a fragmentation warheadand configured for being maneuvered to a predicted interception point(PIP) for intercepting within a probability envelope a flying objecthaving a kill zone of known dimensions, the fragmentation warhead beingconfigured for selectively providing a plurality of fragments directedtowards said probability envelope such that the spacing between any twoadjacent said fragments within the probability envelope is less than atleast one of said known dimensions to ensure that at least one saidfragment impacts said kill zone within said probability envelope,wherein each said fragment is capable of neutralizing the flying objectby impacting said kill zone.
 44. The interception missile according toclaim 43, wherein said kill zone has a length of 0.50 m and a width of0.10 m.
 45. The interception missile according to claim 43 or claim 44,wherein said flying object is any one of: a rocket, a GRAD rocket; aUAV, a manned air vehicle, cruise missile.
 46. A fragmentation warheadconfigured for being mounted in a carrier vehicle, the warhead having alongitudinal dimension along a longitudinal axis and configured forselectively providing a plurality of fragments directed towards a targetarea spaced at an interception spacing from said warhead to provide afragmentation pattern on the target area including at least one set ofsaid fragments in a spaced linear relationship extending to a lengthdimension greater than said longitudinal dimension, wherein adjacentsaid fragments in each said set at the target area are spaced at arespective fragmentation spacing that is within the dimensions of a killzone of an intended target, wherein each said fragment is capable ofneutralizing the intended target by impacting said kill zone.
 47. Thefragmentation warhead according to claim 46, wherein said kill zone hasa length of 0.50 m and a width of 0.10 m.
 48. The fragmentation warheadaccording to claim 46 or claim 47, wherein said flying object is any oneof: a rocket, a GRAD rocket; a UAV, a manned air vehicle, cruisemissile.